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Title:
AIRCRAFT
Document Type and Number:
WIPO Patent Application WO/2019/212744
Kind Code:
A1
Abstract:
Features for an aircraft having efficiently designed propulsive elements are disclosed. The aircraft includes propellers at or near the wingtips that rotate about a rotation axis to provide thrust and also rotate about an axis that is angled with respect to the rotation axis, such as a lateral axis that extends generally along the wing, in order to change the direction of thrust. The aircraft may be a vertical takeoff and landing (VTOL) aircraft, such as a tail-sitter, or other aircraft.

Inventors:
LOVE, Robert Daniel (Inc.3550 General Atomics Cour, San Diego California, 92121, US)
Application Number:
US2019/027668
Publication Date:
November 07, 2019
Filing Date:
April 16, 2019
Export Citation:
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Assignee:
GENERAL ATOMICS AERONAUTICAL SYSTEMS, INC. (3550 General Atomics Court, San Diego, California, 92121, US)
International Classes:
B64C27/52; B64C29/00
Attorney, Agent or Firm:
MALLON, Joseph J. (2040 MAIN STREET, FOURTEENTH FLOORIRVINE, California, 92614, US)
Download PDF:
Claims:
WHAT IS CLAIMED IS:

1. An aircraft comprising:

a wing extending along a lateral axis between first and second wingtip regions and having a central region located along the wing between the first and second wingtip regions, the lateral axis being perpendicular to a longitudinal axis of the aircraft; a first rotor assembly located within the first wingtip region and comprising a first rotor operatively connected to a first motor, the first motor configured to rotate the first rotor about a first rotation axis, and wherein the first rotor assembly is rotatable about the lateral axis such that the first rotation axis rotates about the lateral axis; and a second rotor assembly located within the second wingtip region and comprising a second rotor operatively connected to a second motor, the second motor configured to rotate the second rotor about a second rotation axis, and wherein the second rotor assembly is rotatable about the lateral axis such that the second rotation axis rotates about the lateral axis.

2. The aircraft of Claim 1, wherein the wing has a wingspan extending along the lateral axis between a first wingtip within the first wingtip region and a second wingtip within the second wingtip region, wherein the first rotor assembly is located outboard of the first wingtip and the second rotor assembly is located outboard of the second wingtip.

3. The aircraft of any of Claims 1-2, wherein the wing comprises a leading edge, the first rotor assembly is configured to rotate such that the first rotor is located forward of the leading edge, and the second rotor assembly is configured to rotate such that the second rotor is located forward of the leading edge.

4. The aircraft of any of Claims 1 -3, wherein the wing comprises a top surface and a bottom surface opposite the top surface, the first rotor assembly is configured to rotate such that the first rotor is located above the top surface, and the second rotor assembly is configured to rotate such that the second rotor is located below the bottom surface.

5. The aircraft of any of Claims 1 -4, wherein the first and second rotor assemblies are configured to rotate such that, during takeoff and landing, the first rotor assembly provides a first thrust force in a first direction and the second rotor assembly provides a second thrust force in a second direction that is opposite to the first direction, wherein the first and second directions are substantially perpendicular to the longitudinal axis.

6. The aircraft of any of Claims 1 -5, wherein the first and second rotor assemblies are configured to rotate such that, during flight, the first and second rotor assemblies provide a thrust force in a first direction that is substantially in a same direction as a thrust force provided by a propulsion system.

7. The aircraft of any of Claims 1 -6, further comprising a main rotor coupled with the central region of the wing and configured to rotate about an axis parallel to the longitudinal axis to thereby provide thrust forces to the aircraft in forward flight.

8. The aircraft of Claim 7, wherein the axis of rotation of the main rotor is the longitudinal axis of the aircraft.

9. The aircraft of Claim 7, comprising two or more of the main rotors each rotating about respective axes that are parallel to the longitudinal axis.

10. The aircraft of Claim 7, wherein the wing has a wingspan extending along the lateral axis between a first wingtip within the first wingtip region and a second wingtip within the second wingtip region, and wherein the main rotor comprises at least one blade having a blade length that is at least twenty percent of the wingspan length.

11. The aircraft of Claim 10, wherein the at least one blade has a blade length that is at least fifty percent of the wingspan length.

12. The aircraft of any of Claims 1-10, the wing having an aspect ratio of at least four.

13. The aircraft of Claim 12, the wing having an aspect ratio of at least eight.

14. The aircraft of any of Claims 1-12, wherein the first and second rotor assemblies are configured to rotate to provide thrust forces to the aircraft in forward flight such that an effective aspect ratio of the wing due to the thrust forces is at least one quarter (0.25) greater than an aspect ratio of the wing.

15. The aircraft of any of Claims 1-14, wherein the aircraft is a tail-sitter aircraft.

16. The aircraft of any of Claims 1-15, further comprising a support structure extending rearward from the wing, the aircraft configured to rest on the support structure before takeoff and after landing with the longitudinal axis in a substantially vertical orientation.

17. The aircraft of any of Claims 7-16, further comprising a propulsion system.

18. The aircraft of any of Claims 7-17, wherein the main rotor is coupled with the central region of the wing and operatively connected to the propulsion system, the propulsion system configured to rotate the main rotor about the longitudinal axis of the aircraft.

19. The aircraft of any of Claims 17-18, wherein the propulsion system is an internal combustion system.

20. The aircraft of Claim 19, wherein the propulsion system comprises a turboshaft engine.

21. The aircraft of any of Claims 17-18, wherein the propulsion system is a hybrid electric propulsion system.

22. The aircraft of Claim 21, wherein the propulsion system is a hybrid electric propulsion system comprising:

an electrical power subsystem coupled with the main rotor and configured to supply electrical power to rotate the main rotor;

an electrical energy store coupled with the electrical power subsystem and configured to provide electrical energy to the electrical power subsystem; and

a prime power subsystem coupled with the electrical power subsystem and configured to supply prime power to the electrical power subsystem.

23. The aircraft of any of Claims 1 -22, wherein the first motor and the second motor are electric motors.

24. An aircraft comprising:

a first wing having a first wingtip region and a second wing having a second wingtip region;

a first rotor assembly located within the first wingtip region and comprising a first rotor configured to rotate about a first axis to produce a first thrust, and wherein the first rotor assembly is rotatable about a second axis that is angled with respect to the first axis such that the first axis rotates about the second axis to thereby change a direction of the first thrust; and

a second rotor assembly located within the second wingtip region and comprising a second rotor configured to rotate about a third axis to produce a second thrust, and wherein the second rotor assembly is rotatable about a fourth axis that is angled with respect to the third axis such that the third axis rotates about the fourth axis to thereby change a direction of the second thrust.

25. The aircraft of Claim 24, wherein the first and second rotor assemblies each comprise a nacelle, and wherein each respective nacelle is rotatable about respective second and fourth axes.

26. The aircraft of Claim 24, wherein the first and second rotor assemblies each comprise a body having a fixed portion and a rotatable portion, and wherein each respective rotatable portion is rotatable about respective second and fourth axes.

27. The aircraft of Claim 26, wherein the rotatable portion is a nose cone.

28. The aircraft of Claim 26, wherein the fixed portion is a nose cone.

29. The aircraft of Claim 24, wherein the first and second rotor assemblies each comprise an electric motor configured to rotate respectively the first and second rotors to produce the first and second thrusts.

30. The aircraft of Claim 29, wherein the electric motors are configured to rotate with the respective rotor assembly about respective second and fourth axes.

31. The aircraft of Claim 24, wherein the first rotor assembly is rotatable such that the first thrust contributes to a torque acting on the aircraft, and the second rotor assembly is rotatable such that the second thrust contributes to the torque.

32. The aircraft of Claim 31, wherein the torque is about a longitudinal axis of the aircraft.

33. The aircraft of Claim 24, further comprising a main rotor located inboard of the first and second wingtip regions.

34. The aircraft of Claim 33, wherein the main rotor is configured to create a first torque acting on the aircraft about the longitudinal axis, the first rotor assembly is rotatable such that the first thrust contributes to a second torque acting on the aircraft, the second rotor assembly is rotatable such that the second thrust contributes to the second torque, wherein the second torque is in an opposite direction to the first torque.

35. The aircraft of Claim 34, further comprising multiple main rotors, wherein the first torque is a net torque created by the multiple main rotors.

36. The aircraft of Claim 34, wherein a magnitude of the second torque is approximately equal to a magnitude of the first torque.

37. The aircraft of Claim 24, wherein the first and second rotor assemblies are rotatable such that the first thrust includes a first vector component, the second thrust includes a second vector component, and wherein the first and second vector components are in opposite directions.

38. The aircraft of Claim 24, wherein the aircraft is a vertical takeoff and landing (VTOL) aircraft.

39. The aircraft of Claim 24, wherein the aircraft is a tail-sitter.

40. The aircraft of Claim 24, wherein the first wing extends along the second axis and the second wing extends along the fourth axis.

41. The aircraft of Claim 24, wherein the second and fourth axes are parallel.

42. The aircraft of Claim 24, wherein the second axis is angled with respect to the fourth axis.

43. The aircraft of Claim 24, further comprising:

a third wing having a third wingtip region;

a fourth wing having a fourth wingtip region;

a third rotor assembly located within the third wingtip region and comprising a third rotor configured to rotate about a fifth axis to produce a third thrust, and wherein the third rotor assembly is rotatable about a sixth axis that is angled with respect to the fifth axis such that the fifth axis rotates about the sixth axis to thereby change a direction of the third thrust; and

a fourth rotor assembly located within the fourth wingtip region and comprising a fourth rotor configured to rotate about a seventh axis to produce a fourth thrust, and wherein the fourth rotor assembly is rotatable about an eighth axis that is angled with respect to the seventh axis such that the seventh axis rotates about the eighth axis to thereby change a direction of the fourth thrust.

44. The aircraft of Claim 43, wherein the aircraft has an X-wing configuration.

45. The aircraft of Claim 24, further comprising one or more inboard rotors.

46. The aircraft of Claim 24, further comprising one or more rotor assemblies located within the first and second wing.

47. The aircraft of Claim 24 in combination with any of Claims 1-23.

Description:
AIRCRAFT

INCORPORATION BY REFERENCE TO ANY RELATED APPLICATIONS

[0001] This application claims priority to U.S. provisional application number 62/667,301, entitled AIRCRAFT, and filed May 4, 2018, the disclosure of which is hereby incorporated by reference herein in its entirety for all purposes and forms a part of this specification.

BACKGROUND

Field

[0002] This development relates to aircraft. In particular, propulsive features for aircraft are described.

Description of the Related Art

[0003] Aircraft have different requirements based on the intended type of flight with the aircraft. For example, vertical takeoff and landing (VTOL) aircraft have different requirements for the lifting/landing phases and horizontal flight phases. Conventional approaches to VTOL aircraft are suboptimal with regard to providing sufficient endurance and efficient VTOL capabilities. Other types of aircraft have similar suboptimal designs. Therefore, aircraft features that overcome these drawbacks are desired.

SUMMARY

[0004] The embodiments disclosed herein each have several aspects no single one of which is solely responsible for the disclosure’s desirable attributes. Without limiting the scope of this disclosure, its more prominent features will now be briefly discussed. After considering this discussion, and particularly after reading the section entitled“Detailed Description” one will understand how the features of the embodiments described herein provide advantages over existing approaches to aircraft.

[0005] Described herein are embodiments of systems, devices and methods for an aircraft and systems, devices and methods for features for an aircraft. The aircraft has features that provide for optimal design with regard to particular requirements for a given type of aircraft and/or type of flight. For example, some embodiments are VTOL aircraft that provide for extended endurance flight with efficient VTOL capabilities, among other advantages. As a further example, a tail-sitter takes off from a vertical position sitting upright on the tail, rotates horizontally for forward flight, and then rotates back to vertical for landing on the tail. Some embodiments are tail-sitter aircraft that provide for long endurance flight with efficient VTOL capabilities. Still other types of aircraft and uses may implement the various features described herein.

[0006] Some embodiments of the aircraft include propellers at or near the wingtips. The propellers may rotate about a rotation axis to provide thrust. The propellers may also rotate about a lateral axis that extends generally along the wing in order to change the direction of thrust. The propellers may have a first rotational orientation where thrust is provided parallel to a longitudinal axis of the aircraft, such as a roll axis. The propellers may have a second rotational orientation where thrust is provided perpendicular to the longitudinal axis of the aircraft. For example, the wingtip propellers may be used with VTOL aircraft and in different rotational orientations for different phases of flight, such as providing thrust vectors in different directions to provide counter-torque forces on the aircraft during vertical takeoff or landing, or in the same direction to provide thrust for the aircraft during horizontal flight. The wingtip propellers may be powered by electric motors. The electric motors may rotate with the propellers about the lateral axis.

[0007] The wingtip propellers may be used with a variety of aircraft and have a number of desirable features. For example, the wingtip propellers may provide in a single system 1) the capability to provide reduced drag due to lift (or“induced drag”) during forward flight by creating an up-wash at the wingtips, 2) the capability to provide anti-rotation forces on the aircraft by pointing the thrust vectors in opposite directions (or have a component of each thrust vector that acts in opposite directions) to provide a torque in a first rotation direction that may be opposite to a torque in a second direction that is generated by other spinning features such as rotors or other propellers, etc., 3) the capability to provide or augment thrust forces, for both forward flight of the aircraft by pointing the thrust vectors in the same direction (e.g. forward or rearward) and for vertical flight, for example with the rotors angled during vertical flight to have a vertical component to the thrust, such as angled less than one hundred twenty degrees, less than ninety degree, etc., 4) an increased effective aspect ratio of the wing due to the airflow produced at the wingtips thus allowing for smaller aspect ratio wing structures thereby saving weight, allowing for heavier useful loads, etc., 5) providing attitude control for the aircraft in vertical and forward flight by providing varying levels of thrust at varying angles of rotation of the rotor assemblies, and other capabilities as described herein and that will be apparent to those of ordinary skill in the art.

[0008] Some embodiments of the aircraft can include one or more main rotors. For example, a main rotor with a blade having a larger span than blades of the wingtip propellers may be used for providing thrust for vertical and/or forward flight. Some embodiments may include a single large rotor centrally located on the front of the aircraft with two wingtip propellers. Some embodiments may include two wingtip propellers as well as two or more rotors or propellers located inboard of the wingtips. Some embodiments may include two or more wingtip propellers, as well as one or more other inboard propulsion means, such as one or more rotors, propellers, jet engines, turbojets, turbofans, rockets, ramjets, pulse jets, other engine types, or combinations thereof.

[0009] Some embodiments of the aircraft may include a hybrid electric propulsion system for rotating propellers to generate thrust. A prime power subsystem, such as diesel etc., is coupled with an electric power source and storage, such as an electric generator and battery. The components of the hybrid electric propulsion system are selectively used to generate thrust by rotating blades. Selection of the components of the hybrid electric propulsion system is based on various parameters, such as required power, flight regime (takeoff, landing, horizontal flight, etc.), and others. The hybrid electric propulsion system may power the main source of thrust, such as the main larger rotor, and electric motors may power the wingtip propellers. Such arrangements simplify the configuration, for example, by not having to extend a shaft and associated parts from the main propulsion system to the wingtip motor assemblies, by allowing for the electric motors and propellers to rotate as a unit, as well as other advantages as described herein.

[0010] Another embodiment of the aircraft includes shape-changing aircraft blades. The blades change shape based on various parameters of the system, such as required power, flight regime (takeoff, landing, horizontal flight, etc.), and others. The blades are selectively induced to controllably change shape using piezo elements, which may be piezoelectric. The change in shape alters the span- wise distribution of twist in the blade, i.e. a local pitch along the span of the blade, and thus the amount of thrust generated for a given speed of rotation of the blade. In some embodiments, the aircraft incorporates either the hybrid electric propulsion system or the shape-changing blades. In some embodiments, the aircraft incorporates both the hybrid electric propulsion system and the shape-changing blades.

[0011] In one aspect, an aircraft is described. The aircraft comprises a first wing, a second wing, a first rotor assembly and a second rotor assembly. The first wing has a first wingtip region and the second wing has a second wingtip region. The first rotor assembly is located within the first wingtip region and comprises a first rotor configured to rotate about a first axis to produce a first thrust, and where the first rotor assembly is rotatable about a second axis that is angled with respect to the first axis such that the first axis rotates about the second axis to thereby change a direction of the first thrust. The second rotor assembly is located within the second wingtip region and comprises a second rotor configured to rotate about a third axis to produce a second thrust, and where the second rotor assembly is rotatable about a fourth axis that is angled with respect to the third axis such that the third axis rotates about the fourth axis to thereby change a direction of the second thrust.

[0012] Various embodiments of the various aspects may be implemented. The first and second rotor assemblies may each comprise a nacelle, and each respective nacelle may be rotatable about respective second and fourth axes. The first and second rotor assemblies may each comprise a body having a fixed portion and a rotatable portion, and where each respective rotatable portion is rotatable about respective second and fourth axes. The rotatable portion may be a nose cone. The fixed portion may be a nacelle. The first and second rotor assemblies may each comprise an electric motor configured to rotate respectively the first and second rotors to produce the first and second thrusts. The electric motors may be configured to rotate with the respective rotor assembly about respective second and fourth axes.

[0013] In some embodiments, the first rotor assembly may be rotatable such that the first thrust contributes to a torque acting on the aircraft, and the second rotor assembly may be rotatable such that the second thrust contributes to the torque. The torque may be about a longitudinal axis of the aircraft. The aircraft may further comprise one or more main rotors located inboard of the first and second wingtip regions. The main rotor may be configured to create a first torque acting on the aircraft about the longitudinal axis, the first rotor assembly may be rotatable such that the first thrust contributes to a second torque acting on the aircraft, the second rotor assembly may be rotatable such that the second thrust contributes to the second torque, and the second torque may be in an opposite direction to the first torque. The aircraft may comprise multiple main rotors, where the first torque is a net torque created by the multiple main rotors. A magnitude of the second torque may be approximately equal to a magnitude of the first torque. The first and second rotor assemblies may be rotatable such that the first thrust includes a first vector component, the second thrust includes a second vector component, and the first and second vector components are in opposite directions. The aircraft may be a vertical takeoff and landing (VTOL) aircraft. The aircraft may be a tail-sitter. The first wing may extend along the second axis and the second wing may extend along the fourth axis. The second and fourth axes may be parallel. The second axis may be angled with respect to the fourth axis.

[0014] In some embodiments, the aircraft may further comprise a third wing, a fourth wing, a third rotor assembly and a fourth rotor assembly. The third wing may have a third wingtip region. The fourth wing may have a fourth wingtip region. The third rotor assembly may be located within the third wingtip region and comprise a third rotor configured to rotate about a fifth axis to produce a third thrust, and the third rotor assembly may be rotatable about a sixth axis that is angled with respect to the fifth axis such that the fifth axis rotates about the sixth axis to thereby change a direction of the third thrust. The fourth rotor assembly may be located within the fourth wingtip region and comprise a fourth rotor configured to rotate about a seventh axis to produce a fourth thrust, and the fourth rotor assembly may be rotatable about an eighth axis that is angled with respect to the seventh axis such that the seventh axis rotates about the eighth axis to thereby change a direction of the fourth thrust. The aircraft may have an X-wing configuration. The aircraft may further comprise one or more inboard rotors. The aircraft may further comprise one or more rotor assemblies located within the first and second wing.

[0015] In another aspect, an aircraft is described. The aircraft comprises a wing, a first rotor assembly and a second rotor assembly. The wing extends along a lateral axis between first and second wingtip regions and has a central region located along the wing between the first and second wingtip regions. The first rotor assembly is located within the first wingtip region and comprises a first rotor operatively connected to a first motor. The first motor is configured to rotate the first rotor about a first rotation axis, and the first rotor assembly is rotatable about the lateral axis such that the first rotation axis rotates about the lateral axis. The second rotor assembly is located within the second wingtip region and comprises a second rotor operatively connected to a second motor. The second motor is configured to rotate the second rotor about a second rotation axis, and the second rotor assembly is rotatable about the lateral axis such that the second rotation axis rotates about the lateral axis.

[0016] Various embodiments of the various aspects may be implemented. The first motor and the second motor may be electric motors. The wing may have a wingspan extending along the lateral axis between a first wingtip within the first wingtip region and a second wingtip within the second wingtip region, with the first rotor assembly located outboard of the first wingtip and the second rotor assembly located outboard of the second wingtip. The wing may comprise a leading edge, the first rotor assembly may be configured to rotate such that the first rotor is located forward of the leading edge, and the second rotor assembly may be configured to rotate such that the second rotor is located forward of the leading edge. The wing may comprise a top surface and a bottom surface opposite the top surface, the first rotor assembly may be configured to rotate such that the first rotor is located above of the top surface, and the second rotor assembly may be configured to rotate such that the second rotor is located below of the bottom surface.

[0017] The first and second rotor assemblies may be configured to rotate such that, during takeoff and landing, the first rotor assembly provides a first thrust force in a first direction and the second rotor assembly provides a second thrust force in a second direction that is opposite to the first direction, where the first and second directions are substantially perpendicular to the longitudinal axis. The first and second rotor assemblies may be configured to rotate such that, during flight, the first and second rotor assemblies provide a thrust force in a first direction that is substantially in a same direction as a thrust force provided by a propulsion system.

[0018] The aircraft may have a main rotor coupled with the central region of the wing and configured to rotate about an axis of rotation that is parallel to the longitudinal axis to thereby provide thrust forces to the aircraft in forward flight. The axis of rotation of the main rotor may be the longitudinal axis of the aircraft. There may be two or more of the main rotors each rotating about respective axes that are parallel to the longitudinal axis. The wing may have a wingspan extending along the lateral axis between a first wingtip within the first wingtip region and a second wingtip within the second wingtip region. The main rotor may comprise at least one blade having a blade length that is at least twenty percent of the wingspan. The at least one blade may have a blade length that is at least fifty percent of the wingspan length. The wing may have an aspect ratio of at least four or at least eight. The first and second rotor assemblies may be configured to rotate to provide thrust forces to the aircraft in forward flight such that an effective aspect ratio of the wing due to the thrust forces is at least one quarter (.025) greater than an aspect ratio of the wing.

[0019] The aircraft may be a tail-sitter aircraft. The aircraft may further comprise a support structure extending rearward from the wing, with the aircraft configured to rest on the support structure before takeoff and after landing with a longitudinal axis that is perpendicular to the lateral axis in a substantially vertical orientation.

[0020] The aircraft may further comprise a propulsion system. The aircraft may further comprise a main rotor coupled with the central region of the wing and operatively connected to the propulsion system, with the propulsion system configured to rotate the main rotor about a longitudinal axis of the aircraft. The propulsion system may be an internal combustion system. The propulsion system may comprise a turboshaft engine. The propulsion system may be a hybrid electric propulsion system.

[0021] The propulsion system may be an internal combustion system. The propulsion system may comprise a turboshaft. The propulsion system may be a hybrid electric propulsion system comprising an electrical power subsystem, an electrical energy store, and a prime power subsystem. The electrical power subsystem may be coupled with a main rotor and configured to supply electrical power to rotate the main rotor. The electrical energy store may be coupled with the electrical power subsystem and configured to provide electrical energy to the electrical power subsystem. The prime power subsystem may be coupled with the electrical power subsystem and configured to supply prime power to the electrical power subsystem.

[0022] In another aspect, a hybrid electric propulsion system for an aircraft, such as a tail-sitter, VTOL or other aircraft, having one or more propellers and outboard rotor assemblies with a rotor and an electric motor that are rotatable about a lateral axis, is described. The hybrid electric propulsion system comprises an electrical power subsystem, an electrical energy store and a prime power subsystem. The propeller is configured to provide vertical lift to the aircraft during vertical takeoff and vertical landing phases and to provide horizontal thrust to the aircraft during a horizontal flight phase. The electrical power subsystem is coupled with the propeller and is configured to supply increased electrical power to rotate the propeller at a first speed during the vertical takeoff and vertical landing phases and to supply reduced electrical power to rotate the propeller at a second speed during the horizontal flight phase, where the first speed is greater than the second speed. The electrical energy store is coupled with the electrical power subsystem and is configured to provide electrical energy to the electrical power subsystem during the vertical takeoff and landing phases and to store electrical energy produced by the electrical power subsystem during the horizontal flight phase. The prime power subsystem is coupled with the electrical power subsystem and is configured to supply increased prime power to the electrical power subsystem during the vertical takeoff and vertical landing phases and to supply reduced prime power to the electrical power subsystem during the horizontal flight phase.

[0023] In some embodiments, the electrical power subsystem comprises a generator coupled with the prime power subsystem and an electric motor coupled with the generator and with the propeller. The prime power subsystem may be configured to provide prime power to the generator for production of increased electrical power, and the generator may be configured to supply the increased electrical power to the electric motor to rotate the propeller at high speed during the vertical takeoff and vertical landing phases. The electrical energy store may be coupled with the electric motor, and the electrical energy store may be configured to provide the increased electrical energy to the electric motor during the vertical takeoff and landing phases. The electrical energy store may be coupled with the generator, and the electrical energy store may be configured to store electrical energy produced by the generator during the horizontal flight phase.

[0024] In some embodiments, the prime power subsystem is an internal combustion engine. In some embodiments, the electrical power subsystem supplies a peak electrical power during the vertical takeoff and vertical landing phases. In some embodiments, the prime power subsystem supplies a peak prime power during the vertical takeoff and vertical landing phases. In some embodiments, the electrical and prime power subsystems are configured to collectively produce a first total output of power for liftoff that is at least two times a second total output of power produced for horizontal flight. [0025] In some embodiments, the electrical power subsystem comprises a generator coupled with the prime power subsystem and an electric motor coupled with the generator and with the propeller. The prime power subsystem is configured to provide prime power to the generator for production of increased electrical power, and the generator is configured to supply the increased electrical power to the electric motor to rotate the propeller at the first speed during the vertical takeoff and vertical landing phases. The electrical energy store is coupled with the electric motor, and the electrical energy store is configured to provide the increased electrical energy to the electric motor during the vertical takeoff and landing phases. The electrical energy store is coupled with the generator, and the electrical energy store is configured to store electrical energy produced by the generator during the horizontal flight phase. The electrical power subsystem supplies a peak electrical power during the vertical takeoff and vertical landing phases. The prime power subsystem supplies a peak prime power during the vertical takeoff and vertical landing phases. The electrical and prime power subsystems are configured to collectively produce a first total output of power for liftoff that is at least two times a second total output of power produced for horizontal flight.

[0026] In some embodiments, the propeller comprises a piezo element configured to receive an electric current to change the shape of a propeller blade based on the phase of flight. An increased twist of the blade may be induced by the piezo element for horizontal flight relative to takeoff and landing.

[0027] In another aspect, a tail-sitter aircraft is described. The tail-sitter aircraft comprises a fuselage, a wing, a main rotor, wingtip rotor assemblies, and a hybrid propulsion system. The wingtip rotor assemblies each comprise a rotor and electric motor that are rotatable about a lateral axis of the aircraft to provide counter-torque forces to the aircraft during takeoff and landing. The fuselage has a nose end and a tail end, and the aircraft is configured to be oriented on the ground with the nose end pointing away from the ground. The wing is coupled with the fuselage and is configured to provide lift during a horizontal flight phase. The hybrid propulsion system comprises an electrical power subsystem, an electrical energy store and a prime power subsystem. The electrical power subsystem is coupled with the propeller (of the main rotor and/or the wingtip rotor assemblies) and is configured to supply increased electrical power during vertical takeoff and vertical landing phases and to supply reduced electrical power during the horizontal flight phase. The electrical energy store is coupled with the electrical power subsystem and is configured to provide electrical energy to the electrical power subsystem and to store electrical energy produced by the electrical power subsystem. The prime power subsystem is coupled with the electrical power subsystem and is configured to supply increased prime power during the vertical takeoff and vertical landing phases and to supply reduced prime power during the horizontal flight phase.

[0028] In some embodiments, the propeller (of the main rotor and/or the wingtip rotor assemblies) is configured to provide vertical lift to the aircraft during the vertical takeoff and vertical landing phases and to provide horizontal thrust to the aircraft during a horizontal flight phase. In some embodiments, the electrical and prime power subsystems are configured to collectively rotate the propeller at a relatively higher speed during the vertical takeoff and vertical landing phases and to collectively rotate the propeller at a relatively lower speed during the horizontal flight phase. In some embodiments the electrical energy store provides electrical energy during the vertical takeoff and landing phases and stores electrical energy during the horizontal flight phase. In some embodiments, the electrical power subsystem comprises a generator coupled with the prime power subsystem and an electric motor coupled with the generator and with the propeller, where the prime power subsystem is configured to provide prime power to the generator for production of increased electrical power, and where the generator is configured to supply the increased electrical power to the electric motor to rotate the propeller at high speed during the vertical takeoff and vertical landing phases.

[0029] In some embodiments, the electrical energy store is coupled with the electric motor, and the electrical energy store is configured to provide increased electrical energy to the electric motor during the vertical takeoff and landing phases. In some embodiments, the electrical energy store is coupled with the generator, and the electrical energy store is configured to store electrical energy produced by the generator during the horizontal flight phase. In some embodiments, the electrical power subsystem supplies a peak electrical power during the vertical takeoff and vertical landing phases. In some embodiments, the prime power subsystem supplies a peak prime power during the vertical takeoff and vertical landing phases. In some embodiments, the electrical and prime power subsystems are configured to collectively produce a first total output of power for liftoff that is at least two times a second total output of power produced for horizontal flight. [0030] In another aspect, a method of control for a tail-sitter aircraft is described. The method comprises supplying a first and second prime power from a prime power subsystem to an aircraft engine during, respectively, takeoff/landing and horizontal flight, and supplying a first and second electric power from an electric power source to the aircraft engine during, respectively, takeoff/landing and horizontal flight. A first sum equal to the sum of the first prime and electric powers is greater than a second sum equal to the sum of the second prime and electric powers. The first sum is sufficient to provide vertical lift in an amount at least equal to a force due to gravity on the aircraft, and the second sum is sufficient to sustain horizontal flight. In some embodiments, the first sum is at least two times larger than the second sum. The first sum may be about 300 horsepower. The second sum may be about 60 horsepower.

[0031] In some embodiments, the method further comprises changing the shape of a propeller blade of the aircraft to a first twist for takeoff and landing, and changing the shape of the propeller blade to a second twist for horizontal flight, where the second twist is greater than the first twist. In some embodiments, changing the shape of the propeller blade comprises supplying a current to a piezo element coupled with the blade.

[0032] These and other embodiments are described in greater detail below.

BRIEF DESCRIPTION OF THE DRAWINGS

[0033] The foregoing and other features of the present disclosure will become more fully apparent from the following description and appended claims, taken in conjunction with the accompanying drawings. Understanding that these drawings depict only several embodiments in accordance with the disclosure and are not to be considered limiting of its scope, the disclosure will be described with additional specificity and detail through use of the accompanying drawings. In the following detailed description, reference is made to the accompanying drawings, which form a part hereof. In the drawings, similar symbols typically identify similar components, unless context dictates otherwise. The illustrative embodiments described in the detailed description, drawings, and claims are not meant to be limiting. Other embodiments may be utilized, and other changes may be made, without departing from the spirit or scope of the subject matter presented here. It will be readily understood that the aspects of the present disclosure, as generally described herein, and illustrated in the Figures, can be arranged, substituted, combined, and designed in a wide variety of different configurations, all of which are explicitly contemplated and made part of this disclosure.

[0034] FIG. 1A is a top perspective view of an embodiment of a tail-sitter aircraft having a single main rotor and rotatable wingtip propellers oriented to provide thrust in forward flight.

[0035] FIG. 1B is a top perspective view of the tail-sitter aircraft of FIG. 1A with the rotatable wingtip propellers rotated to provide counter-torque during landing and takeoff.

[0036] FIG. 1C is a top perspective view of another embodiment of a tail-sitter aircraft having the rotatable wingtip propellers of FIGS. 1A-1B and additionally with three main rotors, with the wingtip propellers shown oriented to provide thrust in forward flight.

[0037] FIGS. 1D-1E are schematics of embodiments of propulsion system configurations, having electric and prime power aspects, that may be used with the tail-sitter aircraft of FIGS. 1A-1B.

[0038] FIG. 1F is a table showing an embodiment of control techniques for the various phases of flight that may be used with the aircraft of FIG. 1A-1B.

[0039] FIG. 1G is a table showing an embodiment of aspects of three different schemes for rotating respective wingtip rotor assembly configurations shown in FIGS. 1H-1 J.

[0040] FIGS. 1H-1J are schematics of various embodiments of wingtip rotor assembly configurations that may be used with the various aircraft described herein.

[0041] FIGS. 1K-1L are sequential views of a wingtip rotor assembly shown respectively in a takeoff orientation and a cruise orientation.

[0042] FIG. 1M is a partial schematic of an aircraft showing example thrust vector components and resulting torques that may be created by the various aircraft and systems described herein.

[0043] FIG. 2 is a top perspective view of the tail-sitter aircraft of FIG. 1, with the body of the aircraft shown transparently so that portions of a hybrid propulsion system are visible.

[0044] FIG. 3 is a bottom view of the prime and electric power subsystems of the tail-sitter aircraft of FIG. 1 A, with the body of the aircraft shown transparently so that portions of the prime power subsystem are visible. [0045] FIG. 4 is a schematic of an embodiment of a hybrid propulsion system configuration, having electric and prime power subsystems, that may be used with variants of the tail-sitter aircraft of FIG. 1C.

[0046] FIG. 5 is a flowchart showing an embodiment of a method for operating the hybrid propulsion system of the tail-sitter aircraft of FIG. 1A for vertical takeoff, horizontal flight, and vertical landing.

[0047] FIG. 6A is a flowchart showing an embodiment of a method for operating the hybrid propulsion system of the tail-sitter aircraft of FIG. 1A that may be used for the vertical takeoff and landing phases of flight.

[0048] FIG. 6B is a flowchart showing an embodiment of a method for operating the hybrid propulsion system of the tail-sitter aircraft of FIG. 1A that may be used for the horizontal flight phase.

[0049] FIGS. 7A-7C are perspective views of an embodiment of shape-changing aircraft blades in a first configuration having a first twist that may be used with the tail-sitter aircraft of FIG. 1A for the horizontal flight phase.

[0050] FIGS. 7D-7F are perspective views of the shape-changing aircraft blades of FIGS. 7A-7C in a second configuration having a second twist that may be used with the tail- sitter aircraft of FIG. 1A for the vertical takeoff and landing phases of flight.

[0051] FIG. 8 is a flowchart showing an embodiment of a method for operating the shape-changing aircraft blades of the tail-sitter aircraft of FIG. 1A in flight.

[0052] FIG. 9 is a flowchart showing an embodiment of a method for operating the hybrid propulsion system and the shape-changing aircraft blades of the tail-sitter aircraft of FIG. 1 A in flight.

[0053] FIGS. 10A-10C are various views of another embodiment of a tail-sitter aircraft having a hybrid propulsion system, comprising electric and prime power subsystems, and shape-changing aircraft blades with piezo elements.

[0054] FIGS. 11A-11C are various views of another embodiment of an aircraft having rotatable wingtip rotor assemblies and an X-wing configuration.

[0055] FIGS. 12A-12C are various views of another embodiment of an X-wing aircraft having rotatable wingtip rotor assemblies and fixed inboard rotor assemblies. [0056] FIG. 13 is a perspective view of another embodiment of an aircraft having rotatable wingtip rotor assemblies and fixed inboard rotor assemblies.

[0057] FIG. 14A is a flowchart showing an embodiment of a method for a VTOL aircraft to takeoff using rotatable wingtip rotor assemblies.

[0058] FIG. 14B is a flowchart showing an embodiment of a method for a VTOL aircraft to land using rotatable wingtip rotor assemblies.

DETAILED DESCRIPTION

[0059] The following detailed description is directed to certain specific embodiments of the development. Reference in this specification to“one embodiment,”“an embodiment,” or“in some embodiments” means that a particular feature, structure, or characteristic described in connection with the embodiment is included in at least one embodiment of the invention. The appearances of the phrases“one embodiment,”“an embodiment,” or“in some embodiments” in various places in the specification are not necessarily all referring to the same embodiment, nor are separate or alternative embodiments necessarily mutually exclusive of other embodiments. Moreover, various features are described which may be exhibited by some embodiments and not by others. Similarly, various requirements are described which may be requirements for some embodiments but may not be requirements for other embodiments.

[0060] Aircraft design requires tradeoffs among many parameters, such as drag, weight, endurance, altitude, and others. The rotatable wingtip rotors or propellers described herein provide numerous advantages for various parameters from just a single system. The rotatable wingtip propeller systems can be configured to provide counter-torque forces during takeoff and landing, enhance thrust forces during forward flight, reduce drag, provide attitude control of the aircraft in vertical and forward flight, and/or increase effective aspect ratio of the wings, among other advantages described herein. Various types of aircraft may benefit from the use of the rotatable wingtip propeller systems. Vertical -takeoff and landing (VTOL), such as tail-sitters or forward-facing VTOL aircraft, and others may implement the rotatable wingtip propeller systems described herein and benefit from the many advantages of such systems. The aircraft having rotatable wingtip rotor assemblies described herein may include other features as described, for example, in U.S. patent application no. 15/792,490, filed October 24, 2017, and titled“Tail-Sitter Aircraft with Hybrid Propulsion,” and in U.S. patent application no. 15/792,542, filed October 24, 2017, and titled“Shape Changing Aircraft Blade,” the contents of each of which is incorporated herein by reference in its entirety for all purposes and forms a part of this specification.

[0061] FIG. 1A is a top perspective view of an embodiment of a tail-sitter aircraft 100 that includes wingtip rotor assemblies 230. As one example of an aircraft that can benefit from the rotatable wingtip rotor assemblies 230, VTOL aircraft have challenging requirements, including power and lift requirements. The rotatable wingtip rotor assemblies 230 can be powered in such aircraft using various types of power sources. For example, electric motors may be used, thus simplifying the main propulsion system design, for example by not needing shafts extending to the wingtips, allowing for simpler rotation of the wingtip propeller systems, and other. The main propulsion system can also power the power source for the rotor assemblies, such as the electric motors. In some embodiments, there may be dual main motors with partial conversion of one of the main motor’s shaft power electrical energy, for example during cruise.

[0062] The propulsion system 300 may be an internal combustion type system. The system 300 may comprise a gas turbine configured to produce shaft power. The system 300 may rotate the main rotor 210. The system 300 may power a generator to provide electric power to electrical components, such as the electric motors as described herein. The system 300 may comprise a turboshaft.

[0063] In some embodiments, the propulsion system 300 may be a hybrid system. For example, internal combustion engines can be sized to provide the high power demand of the takeoff and landing phases, but that sizing results in unnecessary extra power and weight for the horizontal flight phase. Electric propulsion systems provide high power more efficiently, but have insufficient power capacity to provide the necessary endurance and/or payload lifting capacity. A hybrid propulsion system as described herein is ideal to supply both: (1) high power by combining peak prime power and stored electrical energy for a short period of time for VTOL and (2) reduced power from the small prime power subsystem during the slower, long-duration horizontal or loiter phase. Although hybrid power sources have traditionally been considered more complicated (and therefore less reliable) and also undesirably heavier, the design of the hybrid-propulsion system described herein provides for a lower weight propulsion system than previously used for conventional VTOL tail-sitter aircraft. In addition, the hybrid power system described herein may make the aircraft more reliable, since it can provide power for engine-out emergency landings and enable bursts of speed to increase platform responsiveness.

[0064] Vertical-takeoff and landing (VTOL) aircraft also have challenging lift requirements with respect to propellers. As used herein, unless otherwise indicated explicitly or by context, the terms“propeller,”“blade,” and“rotor” may be used interchangeably. An aircraft blade, such as a propeller, rotor, etc., converts rotary motion from an engine or other mechanical power source, to provide propulsive forces such as lift during vertical takeoff or thrust during horizontal flight. It comprises a rotating power-driven hub, to which several radial airfoil-section blades are attached, such that the whole assembly rotates about a longitudinal axis. The propeller attaches to the power source's driveshaft either directly or through reduction gearing. The twist of the blade defines a series of localized pitches along the span of the blade, where the span extends from an inner hub to the outer tip of the blade. The pitch of a propeller blade refers to turning the angle of attack of the aircraft blades into or out of the wind to control the production or absorption of power. The blade pitch may be described as "coarse" for a coarser angle and "fine" for a finer angle. The blade pitch may be described in units of distance/rotation assuming no slip. Low pitch yields good low speed acceleration (and climb rate in an aircraft) while high pitch optimizes high speed performance and economy. The pitch may be altered by changing the shape of the blade by altering the amount of twist while keeping the root of the blade fixed. The pitch may also be altered by rotating the entire blade along an axis extending along the span of the blade, such as by rotating the blade at the root. This rotation does not change the shape of the blade but rather rotates the entire blade while maintaining the shape. Both twist inducement to change the shape and pitch changes via blade rotation may be simultaneously incorporated.

[0065] A propeller blade's propulsive output, such as lift or thrust, depends on the angle of attack combined with its speed. Because the velocity of a propeller blade varies from the inner end or hub to the outer end or tip, it has a twisted form in order for the propulsive output to remain approximately constant along the length of the blade; this is called washout. Varying the pitch in flight may give optimum thrust over the maximum amount of the aircraft's speed range, from takeoff and climb to cruise. [0066] In some embodiments, the geometries of efficient rotors and propellers may be fundamentally different from one another due to the difference in the angle and speed of incoming air. Embodiments of shape-changing blades described herein facilitate the ability of the blade to behave like an efficient rotor. In addition, the variations in geometry will enable the blade to efficiently address variations in air density as the aircraft changes altitude. Some embodiments of the blade provide a variable-geometry propeller structural configuration, along with a sensing and control method designed to provide efficient transition and flight throughout flight regimes including vertical or horizontal takeoff, horizontal flight, and landing. The shape-changing blade structural configuration described herein also adapts well to varying air density.

[0067] Some embodiments provide a variable-pitch blade which changes twist using actuation generated with the piezoelectric effect to such an extent that it may also act as an efficient rotor which provides cyclic and collective control. Some embodiments provide a novel placement of the piezoelectric material on the blade planform and within the internal structure to amplify the deflections, and thereby the overall twist of the blade. Some embodiments enable a blade shaped initially like a typical propeller, which is typically inefficient during takeoff and landing due to low advance ratios, to perform as efficiently as a helicopter rotor on takeoff and landing. Some embodiments also provide information regarding aeroelastic blade deflections such that the blade geometry can be adjusted to maintain efficient flight in various flight conditions.

[0068] Embodiments of the variable-geometry aircraft blades described herein may be used on VTOL aircraft, such as a“tail-sitter” aircraft that takes off and lands vertically sitting on its tail with its nose pointing upward. Thus, the blades may be used as rotors for takeoff (similar to helicopter) and as a propeller for sustained flight (similar to typical airplane), with the blades changing shape as needed during these flight regimes.

[0069] For ease of description only, an X-Y-Z axis system is shown in FIG 1A. The positive X, Y and Z directions are indicated by the axes as shown. The positive X direction points in the front direction of the aircraft 100, and thus the negative X direction is toward the rear direction of the aircraft 100 which is opposite the front direction. The positive Y direction is perpendicular to the X-axis and points in the left side direction of the aircraft 100, and the negative Y direction is toward the right side direction of the aircraft 100 which is opposite the left direction. The positive Z direction is perpendicular to the X and Y axes and follows the “right hand rule” where the positive Z direction points toward the top side direction of the aircraft 100 and the negative Z direction is toward the bottom side direction of the aircraft 100 which is opposite the top side direction. The longitudinal axis of the aircraft 100 is aligned with the X-axis, as indicated in FIG. 1A. The X-Y-Z axis system defines several geometric planes, for ease of description only. An X-Y plane is a plane that intersects the X and Y axes. An X-Z plane is a plane that intersects the X and Z axes. A Y-Z plane is a plane that intersects the Y and Z axes. Unless otherwise noted,“inner,”“inward,” and the like refer to directions generally toward the longitudinal axis, and“outer,”“outward,” and the like refer to directions generally away from the longitudinal axis.

[0070] As further shown in FIG. 1 A, the tail-sitter aircraft 100 can take off and land vertically with its tail on the ground. In some embodiments, other structures instead of or in addition to the tail may support the aircraft 100 on the ground, such as the wings, ground supports, and/or other structures. The term“tail sitter” includes all of these embodiments.

[0071] The aircraft 100 includes a wing 119. The wing may comprise one or more wing segments. The wing 119 may include a left wing 120 and a right wing 130. The left and right wings 120, 130 have respective leading edges 121, 131 on a forward edge of the wings. The wing 119 may include a center wing 110 as well. The wing 119, such as the center wing 110, attaches to a body 301, such as a central body or fuselage. The body 301 may be short in the longitudinal direction, such that the aircraft 100 is essentially a“flying wing.” A main rotor 210 attached to a front side of the body 301 includes blades 212, 214, 216 that rotate to provide the primary lifting and thrust forces. The propulsion system 105, which may be a hybrid propulsion system as further described herein, powers the main rotor 210.

[0072] In some embodiments, the wing 119 or portions thereof extends along a lateral axis or axes between first and second wingtip regions 122, 132 and has a central region 106 located along the wing 119 between the first and second wingtip regions 122, 132, the lateral axis being perpendicular to a longitudinal axis of the aircraft 100. As further described herein, it is understood that the“lateral axis” as used herein may be parallel to or coincide with the aircraft’s pitch axis (which may be the Y axis), or in some embodiments the lateral axis may be a local rotational axis that is angled with respect to the pitch axis. The lateral axis therefore may or may not extend along an axis of the wing 119, and the embodiment shown in FIG. 1 A is merely one example.

[0073] The aircraft 100 includes two rotatable rotor assemblies 230. One rotor assembly 230 is located at a left wingtip region 122 of the left wing 120 and the second rotor assembly 230 is located at a right wingtip region 132 of the right wing 130. The regions 122, 132 include wingtips 124, 134 respectively. The wingtips 124, 134 are located at the outermost edges of the wings 120, 130 in the +Y and - Y directions. The wingtip regions 122, 132 include the wingtips 124, 134 as well as portions of the wings located inboard (toward the longitudinal axis) from the wingtips 124, 134. The wingtip regions 122, 132 may also include locations outboard (away from the longitudinal axis) from the wingtips 124, 134. In some embodiments, a first rotor assembly 230 is located within the first wingtip region 122 and comprises a first rotor 240 operatively connected to a first motor 233, such as an electric motor, with the first motor 233 configured to rotate the first rotor 240 about a first rotation axis, and the first rotor assembly 230 is rotatable about the lateral axis such that the first rotation axis rotates about the lateral axis. In some embodiments, a second rotor assembly 230 is located within the second wingtip region 132 and comprises a second rotor 240 operatively connected to a second motor 233, such as an electric motor, with the second motor 233 configured to rotate the second rotor 240 about a second rotation axis, and the second rotor assembly 230 is rotatable about the lateral axis such that the second rotation axis rotates about the lateral axis.

[0074] The rotor assemblies 230 may rotate about the lateral axis, which may be oriented as shown or otherwise oriented as discussed. Thus, in some embodiments, the rotor assemblies 230 may each rotate about their own local lateral axes, which may be angled with respect to each other. The rotor assemblies 230 may rotate about a lateral axis that is angled with respect to the X, Y and/or Z directions.

[0075] The rotors 240 may be powered by the motors 233 to provide counter torque during VTOL and provide thrust during cruise. The motors 233 may rotate ninety or approximately ninety degrees during VTOL and rotate back into the cruise position during the transition and conversion process. The existence of two rotors 240 produces the counter-torque without generating a net side force, unlike a helicopter tail rotor where the side force must be trimmed out with the rotor. This eliminated net force enables reduced power required for the counter torque system. The rotors 240 are raised up so that when they pivot, the minimum downforce is imposed on the wing 119.

[0076] At speeds above a certain amount, such as 120 knots, the motors 233 may provide significant thrust increases. In forward flight, the wingtip mounted rotors 240 when rotated to an orientation to provide thrust also provide significant up wash on the wing tips 124, 134. This increases the effective aspect ratio, increasing the wing maximum lift coefficient, increasing the Reynolds number at the tip to increase airfoil efficiency and thereby increasing the overall lift to drag ratio of the wing 119. Such upwash from the wingtip rotor assemblies 230 may provide a similar aerodynamic effect as winglets, which are vertical structures at the wingtips. In addition, the wingtip rotor assemblies 230 help prevent tip stall enabling wing twist customization as needed for best aerodynamic performance. In contrast, propellers that are embedded in the wing tip regions 122, 132 would not be dual use since they would not provide both counter torque and thrust for cruise. In addition, they would have significantly higher disk loading requiring more electrical power conversion (e.g., with electric motors) and they would provide no reduction in induced drag. Further, in forward flight, the wingtip rotor assemblies 230 may be rotated about the lateral axis to orient the thrust vectors farther up or down to thereby provide control of the aircraft. One or both of the rotor assemblies 230 can be rotated in a first direction about the lateral axis to create or increase an upward thrust component (+Z direction) at the respective wingtip or wingtips, and rotated in a second direction about the lateral axis that is opposite to the first direction to create or increase a downward thrust component (-Z direction) at the respective wingtip or wingtips. For example, both rotor assemblies 230 may be rotated an equal amount and in the same direction to cause the aircraft 100 to pitch up or down, i.e. a positive or negative rotation about the Y axis or an axis parallel thereto. For example, one rotor assembly 230 can be rotated a first direction and the other rotor assembly 230 can be rotated a second direction opposite to the first direction to cause the aircraft 100 to roll, i.e. a positive or negative rotation about the X axis or an axis parallel thereto. In some embodiments, one of the rotor assemblies 230 (at various rotational angles about the lateral axis) can be provided more power than, and thus create more thrust than, the other rotor assembly 230 to cause the aircraft to yaw, i.e. a positive or negative rotation about the Z axis. [0077] The aircraft 100 may provide significant electrical power to support payloads. When needed, thrust production can be shifted to the main rotor 210 and powered by the propulsion system 105 engine so that more electrical power is available for the payloads.

[0078] The rotor assemblies 230 each include a body 232, which may be a nacelle. The body 232 attaches to the wings 120, 130. The body 232 may be located outboard from the wingtips 124, 134 as shown. The wing 119 may have a wingspan extending along the lateral axis between the first wingtip 124 within the first wingtip region 122 and the second wingtip 134 within the second wingtip region 132, with the first (left) rotor assembly 230 located outboard of the first wingtip 124 and the second (right) rotor assembly 230 located outboard of the second wingtip 134. In some embodiments, the body 232 may be located at, adjacent or otherwise near the wingtips 124, 134, inboard from the wingtips 124, 134, outboard from the wingtips 124, 134, or combinations thereof. The body 232 includes a fixed portion 234 and a rotatable portion 236. The rotatable portion 236 is located forward of the fixed portion 234. The rotatable portion 236 rotates about a lateral axis, as indicated. The lateral axis may be parallel to the Y-axis, which may be the pitch axis of the aircraft 100. In some embodiments, the entire body 232 may rotate about the lateral axis. In some embodiments, the lateral axis may be angled with respect to the Y-axis and/or X-axis. In some embodiments, each lateral axis at each wingtip region 122, 132 may not be aligned with each other, for example the two lateral axes may be angled with respect to each other.

[0079] The rotor assembly includes a rotor 240 having blades 242. There may be three blades 242 as shown. There may be two, four or more blades 242. The rotor 240 attaches to the rotatable portion 236. A motor 233, such as an electric motor, is operatively connected to the rotor 240. The motor 233 rotates the rotor 240 and blades 242. In some embodiments, the motor 233 may rotate with the rotor 240, for example with the rotatable portion 236. Thus in some embodiments, the entire rotor assembly 230 may rotate relative to the wingtips 124, 134. The rotor assemblies 230 are configured to rotate such that the rotors 240 are located forward of the respective leading edges 121, 131.

[0080] In FIG. 1A, the rotor assemblies 230 are rotated to a forward (+X) configuration. The rotor assemblies 230 may be rotatable about the lateral axis to the orientation shown in FIG. 1A to provide and/or augment thrust to the aircraft 100 in forward flight. The rotor assemblies 230 may be configured to rotate such that, during flight, the rotor assemblies 230 provide a thrust force in a direction that is substantially in a same direction as a thrust force provided by the main rotor 210.

[0081] In FIG. 1B, the rotor assemblies 230 are rotated to a counter- torque configuration. The left rotor assembly 230 is rotated to face the rotor 240 in the -Z direction and the right rotor assembly 230 is rotated to face the rotor 240 in the +Z direction. The left rotor 240 produces a thrust vector pointing in the -Z direction as shown. The right rotor 240 produces a thrust vector pointing in the +Z direction as shown. The rotors 240 may produce thrust vectors aligned with the Z-axis or at an angle to the Z-axis. The rotors 240 may produce thrust vectors that are perpendicular to the longitudinal axis of the aircraft. The rotors 240 may produce thrust vectors that are parallel, or that are not parallel, to each other (whether in the same or opposite directions).

[0082] The rotors 240 may produce thrust vectors or components thereof that contribute to a torque or moment on the aircraft 100. The thrust forces may entirely contribute to such torque. In some embodiments, components of the thrust vectors may contribute to such torque. For example, a thrust vector may be broken down into X, Y and/or Z components, each component having a magnitude and direction. The two thrust vectors produced by respective rotor assemblies 230 may have a vector component or components that are opposite, or generally opposite, in direction to each other. Such component or components may be equal or unequal in magnitude.

[0083] The torque on the aircraft created by the rotor assemblies 230 may be applied about or generally about the longitudinal axis. The torque may be termed an“anti torque” where the torque generated by the rotor assemblies 230 is opposite in direction to a torque on the aircraft produced by other propulsion systems, such as the propulsion system 105 which may include the main rotor 210, or more than one main rotor 210 as described herein. The main rotor 210 and/or other propulsion features may produce a first torque about the longitudinal or other axis in a first rotational direction and the rotor assemblies 230 may produce a counter torque which is a second torque about the same axis but in a second rotational direction that is opposite to the first rotational direction. The counter-torque may be equal, lesser, or greater in magnitude relative to the first torque. The first torque may be a net torque on the aircraft created by multiple main rotors 210, as described herein. [0084] The wing 119 may comprise a top surface 125 and a bottom surface 126 opposite the top surface 125, and the first (left) rotor assembly 230 may be configured to rotate such that the first rotor 240 is located above the top surface 125 and the second (right) rotor assembly 230 may be configured to rotate such that the second rotor 240 is located below the bottom surface 126. In some embodiments, the rotor assemblies 230 may not be directly above or below the respective surfaces 125, 126, for example the rotor assemblies 230 may be forward of the leading edges 121, 131 such that some or all of the rotors 240 are not directly above or below the respective surfaces 125, 126. The rotors 240 may be above or below these or other portions of the top or bottom surfaces 125, 126 that are located within the respective wingtip regions 122, 132. In some embodiments, the first and second rotor assemblies 230 are configured to rotate such that, during takeoff and landing, the first rotor assembly 230 provides a first thrust force in a first direction and the second rotor assembly 230 provides a second thrust force in a second direction that is opposite to the first direction, where the first and second directions are substantially perpendicular to the longitudinal axis.

[0085] In some embodiments, the entire rotor assembly 230 may rotate. For example, the body 232, such as a nacelle, and the rotor 240 may rotate together as a unit. Thus the entire body 232 may be rotatable. The motor 233, such as an electric motor, may rotate with the rotatable portions of the rotor assembly 230. The motor 233 may be located within the rotatable portion 236. In some embodiments, the motor 233 may be located within the body 232 which may entirely rotate as described.

[0086] The various rotors may have various sizes. For example, the wing 119 may have a wingspan, which may extend along the lateral axis, between the first wingtip 124 within the first wingtip region 122 and the second wingtip 134 within the second wingtip region 132. The main rotor 210 may comprise at least one blade 212 having a blade length that is at least ten, twenty, thirty, forty, fifty, sixty, seventy, eighty, ninety percent or more of half the wingspan length. In some embodiments, the main rotor 210 may comprise at least one blade 212 having a blade length that is longer than half the wingspan length. The main rotor 210 may comprise at least one blade 212 that extends laterally out opposite sides from the hub 218 such that the blade is symmetric with respect to the hub 218. Such blade may have a blade length that is at least ten, twenty, thirty, forty, fifty, sixty, seventy, eighty, ninety percent or more of the wingspan length, or that is longer than the wingspan length. [0087] The rotor assemblies 230 may be rotated to other angular orientations. For example, the rotor assemblies 230 may be rotated to angular orientations in between those depicted in FIGS. 1 A and 1B, for example between zero and ninety degrees, between zero and one hundred twenty degrees, between zero and one hundred eighty degrees, or other angular amounts or ranges. The rotor assemblies 230 may increase the thrust in the vertical direction with the rotor assemblies 230 providing thrust at an angle to the longitudinal X direction. For example, if the rotor assemblies 230 pivot to seventy degrees for counter torque, the rotor assemblies 230 would also provide a net lift (and result in reduced downforce on the wing). Other angular orientations may be used. Further, the rotor assemblies 230 may be rotated to provide thrust at an angle to the longitudinal X direction to control the aircraft attitude, as further described herein.

[0088] The wing 119 may have an aspect ratio. The term“aspect ratio” may refer to its usual and customary meaning, including but not limited to the ratio of the wingspan to the mean chord of the wing 119. The aspect ratio may be at least two, three, four, five, six, seven, eight, nine, ten, eleven, twelve, thirteen, fourteen, fifteen, sixteen, seventeen, eighteen, nineteen, twenty, or more. The use of the rotor assemblies 230 for thrust during forward flight may increase the aspect ratio of the wing 119 to provide a greater“effective aspect ratio.” The first and second rotor assemblies 230 may be configured to rotate to provide thrust forces to the aircraft 100 in forward flight such that an effective aspect ratio of the wing 119 due to the thrust forces is at least one tenth (0.1), one eight (0.125), one fifth (0.2), one quarter (0.25), one third (0.33), one half (0.5), three quarters (0.75), one, two, three, four, five or more greater than an aspect ratio of the wing 119. For example, the wing 119 may have an aspect ratio of twelve, and due to aerodynamic effects of the rotor assemblies 230 the wing 119 may have an effective aspect ratio in forward flight of twelve and one quarter, thirteen or fourteen, etc.

[0089] The aircraft 100 may be a tail-sitter. In some embodiments, a support structure such as left and right tails 180, 190, may extend rearward from the wing 119, with the aircraft 100 configured to rest on the support structure before takeoff and after landing with the longitudinal axis in a substantially vertical orientation. Other types of VTOL aircraft, or other types of aircraft, may utilize the rotor assemblies 230.

[0090] The aircraft 100 has the propulsion system 105, which may include the main rotor 210 as shown. Other types of propulsion systems may be used. As shown, the main rotor 210 may be coupled with the central region 106 of the wing 119 and operatively connected to a power subsystem of the propulsion system 105, with the propulsion system 105 configured to rotate the main rotor 105 about the longitudinal axis of the aircraft 100. The propulsion system 105 may be a variety of types of systems, such as the hybrid electric propulsion system as shown. The propulsion system 105 may include an electrical power subsystem 200, an electrical energy store 211, and a prime power subsystem 300. The electrical power subsystem 200 may be coupled with the main rotor 210 and configured to supply electrical power to rotate the main rotor 210. The electrical energy store 211 may be coupled with the electrical power subsystem 200 and configured to provide electrical energy to the electrical power subsystem 200. The prime power subsystem 300 may be coupled with the electrical power subsystem 200 and configured to supply prime power to the electrical power subsystem 200. The electrical and prime power subsystems 200, 300 may be located within the central body 301. In some embodiments, the subsystems may be in different locations. For example, the electrical power subsystem 200 may be in the central body 301 and the prime power subsystem or subsystems 300 may be in one or both of housings 160, 170.

[0091] The propulsion system 105 as shown includes an electrical power subsystem 200 and a prime power subsystem 300, as described in further detail herein. The propulsion system 105 may be entirely a prime, e.g. combustion, power system, or an all electric system. As shown, the hybrid propulsion system 105 includes an electrical power subsystem 200 operatively connected to a prime power subsystem 300 and to the main rotor 210. The main rotor 210 may include blades 212, 214, 216 having features for changing geometric shape, such as the piezo elements as described herein. The prime power subsystem 300 is operatively connected to the electrical power subsystem 200 to provide a hybrid propulsion power source to controllably rotate the main rotor 210. In some embodiments, there may only be one main rotor 210 or more than one main rotor 210, for example, three, five, seven or more (see, e.g., FIG. 1C). In some embodiments, there may be two or more electrical power subsystems 200 and/or two or more prime power subsystems 300.

[0092] The aircraft 100 can fly in vertical and horizontal flight modes. Vertical flight mode is where lift to the aircraft 100 is provided primarily from thrust produced by the rotor 210 (and in some embodiments also the rotors 240) and not primarily from the wings. In vertical flight mode, the longitudinal axis is generally aligned with a geographic vertical or line of action of gravity. Vertical flight mode includes trajectories of the aircraft 100 where the longitudinal axis of the aircraft 100 is not exactly aligned with the geographic vertical or line of action of gravity. Horizontal flight mode is where lift to the aircraft 100 is provided primarily from the wings and not primarily from the rotors. In horizontal flight mode, the Z axes of the aircraft 100 are generally aligned with the geographic vertical or line of action of gravity. Horizontal flight mode includes trajectories of the aircraft 100 where the Z axes of the aircraft 100 are not exactly aligned with the geographic vertical or line of action of gravity. There may be some overlap between the two flight modes where the aircraft 100 is simultaneously in vertical and horizontal flight mode, where lift is provided approximately equally from the propellers and the wings. The aircraft 100 may be transitioning between horizontal and vertical flight modes. Further, the aircraft 100 may be flying in vertical flight mode but in a horizontal direction, for example where the rotors are providing the primary lift but the aircraft 100 is travelling horizontally.

[0093] The main rotor 210 provides vertical lift to the aircraft 100 during the vertical flight phases (e.g. takeoff and landing) and provides horizontal thrust to the aircraft 100 during the horizontal flight phase. The electrical subsystem 200 and prime power subsystem 300 collectively rotate the main rotor 210 at a relatively higher speed during the vertical takeoff and vertical landing phases and collectively rotate the main rotor 210 at a relatively lower speed during the horizontal flight phase. The speed of the main rotor 210 speed is measured in revolutions per minute (RPM). The main rotor 210 may act like a rotor during the vertical flight phases and as a propeller during the horizontal flight phase.

[0094] In horizontal flight mode, the wing 119 provides lift to the aircraft 100. The wing 119 may include the center, left and right wings 110, 120, 130 as shown. In some embodiments, the wing 119 may be one continuous wing without such sections. The wing 119 may have the swept shape as shown and as further described herein. In some embodiments, the wing 129 may have other shapes, such as symmetric, no sweep, swept backward, rectangular, elliptical, trapezoidal, other suitable shapes, or combinations thereof.

[0095] As shown, the center wing 110 is located generally near the center of the aircraft, and may be approximately symmetric with respect to the X-Z plane. The left and right wings 120, 130 are located, respectively, to the left and right sides of the center wing 110. All or some of the center, left and right wings 110, 120, 130 may be formed from composite, metallic, other suitable materials, or combinations thereof. All or some of the center, left and right wings 110, 120, 130 may include outer skin with internal spars and/or other internal structural components. All or some of the center, left and right wings 110, 120, 130 may have a variety of external airfoil shapes that produce lift at particular angles of attack with respect to the freestream flow. In some embodiments the center wing 110 may not have such an airfoil shape, where lift is provided by the right and left wings 120, 130 only.

[0096] The center wing 110 includes a left portion 140 and right portion 150. An inner end of the left portion 140 is connected to a left side of a body 301 of the power subsystems 200, 300, and the opposite, outer end of the left portion 140 is connected to a housing 160 on the left side of the aircraft 100. The body 301 forms part of a“fuselage” of the aircraft 100. The fuselage may or may not include the housings 160, 170 and various connected components described herein. The fuselage maybe longer or shorter than shown. In some embodiments, there may not be a fuselage. The left wing 120 is connected to the left side of the housing 160 that is on the left side of the aircraft 100. An inner end of the right portion 150 is connected to a right side of the body 301 , and the opposite outer end of the right portion 150 is connected to a left side of the housing 170. The right wing 130 is connected to the right side of the housing 170 that is on the right side of the aircraft 100.

[0097] The housings 160, 170 may include features for supporting the aircraft 100 in an upright position, for example before takeoff and after landing. The housings 160, 170 are each, respectively, connected to a left and right tail 180, 190. The tails 180, 190 may be formed from the same or different materials as the housings 160, 170. The tails 180, 190 may be configured to provide takeoff supports on which the aircraft 100 sits before vertical takeoff. The tails 180, 190 may also be configured to provide landing supports when the aircraft 100 lands vertically. The housings 160, 170 and tails 180, 190 may form part of a“fuselage” of the aircraft 100, as described herein. The tails 180, 190 may extend rearwardly generally in a -X direction as shown. In some embodiments, the tails 180, 190 may extend rearwardly at an angle with respect to the -X direction or to the longitudinal axis. There may be only one of the tails 180, 190, for example disposed at or near the central region 106 of the aircraft 100. Other embodiments are shown and described herein, for example with respect to FIGS. 10A- 10C. There may one such centrally-located support structure with supports on both wings, which may be along the length of the wings, at the wingtip regions 122, 132, at the wingtips 124, 134, or located farther inboard.

[0098] As shown, the left tail 180 includes a top vertical fin 182 and a bottom vertical fin 184. The top vertical fin 182 extends in the positive Z direction and is parallel or approximately parallel to the X-Z plane. In some embodiments, the vertical fin 180 may not be parallel or approximately parallel to the X-Z plane. The bottom vertical fin 184 extends in the negative Z direction and is parallel or approximately parallel to the X-Z plane. In some embodiments, the bottom vertical fin 184 may not be parallel or approximately parallel to the X-Z plane. The top and bottom vertical fins 182, 184 provide a larger footprint for the respective tails 180, 190 to facilitate balance of the aircraft 100 during vertical takeoff and landing. In some embodiments, there may be more than one top vertical fin 182 and/or more than one bottom vertical fin 184. For example, there may be multiple angled top and/or bottom vertical fins 182, 184.

[0099] The right tail 190 includes a top vertical fin 192 and a bottom vertical fin 194. The tail 190 may be analogous to the tail 180, where the top vertical fin 192 is analogous to the top vertical fin 182 of the left tail 180, and the bottom vertical fin 194 is analogous to the bottom vertical fin 184 of the tail 180. By“analogous” it is meant the respective components may have the same or similar features and/or functionalities.

[0100] In some embodiments, there may be fewer or more than two tails 180, 190. There may only be a single tail 180 or 190, for example in the center of the aircraft 100. There may be more than two tails 180 or 190, for example if there are more than two electrical power subsystems 200, as described above. Another embodiment showing an alternate configuration for a tail of the aircraft 100 is shown in FIGS. 10A-10C.

[0101] The main rotor 210 includes three blades 212, 214, 216. In some embodiments, the main rotor 210 may have less than or more than three blades, such as two, four, five, six, seven, eight, or more blades. A hub 218 is located at the front side of the main rotor 210 and centrally located with respect to the blades 212, 214, 216. The hub 218 is rotated by the electrical and prime power subsystems 200, 300, as described herein. The hub 218 is connected to inner ends of each of the blades 212, 214, 216. Thus, rotation of the hub 218 rotates the blades 212, 214, 216. Rotation of the blades 212, 214, 216 provides lift during vertical flight mode, such as vertical takeoff and landing, as well as thrust during forward or horizontal flight mode, such as level flight. The blades 212, 214, 216 thus act as rotors, similar to helicopter rotors, during vertical flight mode, and act as propellers, similar to propellers on horizontal takeoff airplanes, during horizontal flight mode.

[0102] In some embodiments the blades 212, 214, 216 and/or 240 may change shape. The wingtip blades 240 may have the same or similar features as the blades 212, 214, 216 described herein. The blades 212, 214, 216 may have a first shape for vertical flight mode, such as during vertical takeoff and landing, and another shape for horizontal flight mode, such as during level flight. The blades 212, 214, 216 may all have the same shape at the same time. In some embodiments, the blades 212, 214, 216 may vary in shape with respect to each other. In some embodiments, the blades 212, 214, 216 connected to the same hub 218 may vary in shape with respect to each other. In some embodiments, the blades 212, 214, 216 connected to the hub 218 on the left may vary in shape with respect to the blades 212, 214, 216 connected to the hub 218 on the right. The blades 212, 214, 216 may change shape using piezo elements, as described herein. Alternatively or in addition, the blades 212, 214, 216 may change shape using other devices, such as torque tubes, etc. In some embodiments, the blades 212, 214, 216 may incorporate other shape changes, such as changes to the shape of the airfoils of the blades 212, 214, 216.

[0103] In some embodiments, the blades may not change shape. For example, the aircraft 100 may include blades 212, 214, 216 that rotate along respective longitudinal axes of the respective blades. Such rotations may uniformly change the pitch of the blades 212, 214, 216 along the longitudinal span of the blades 212, 214, 216. A torque tube or other device may be used to rotate the blades. In some embodiments, the blades 212, 214, 216 may change shape as well as be rotated. In some embodiments, the blades 212, 214, 216 may not change shape or rotate, for example they may be static.

[0104] The change in shape and/or rotation of the blades 212, 214, 216 may be based on required propulsive force, such as thrust or lift, from the blades 212, 214, 216. The thrust force is generated by rotation of the blades 212, 214, 216 and provides movement to the aircraft 100 in the forward or positive X direction. More propulsive force from the blades 212, 214, 216 in the form of lift is required during vertical flight mode. The blades 212, 214, 216 thus may have a first shape and rotational speed that provides more propulsive force during vertical flight mode and a second shape and rotational speed that provides relatively less propulsive force during horizontal (i.e. forward) flight mode. In vertical flight mode, the propulsive force from the blades 212, 214, 216 provides the primary lifting force to the aircraft 100. In horizontal flight mode, the propulsive force from the blades 212, 214, 216 provides thrust such that the wings 120, 130, 140 and/or 150 provide the primary lifting force to the aircraft 100. Further details of the shape-changing aspects of the blades 212, 214, 216 are provided herein, for example with respect to FIGS. 7A-12.

[0105] FIG. 1C is an embodiment of the aircraft 100 having the wingtip rotor assemblies 230 and with three main rotors 210. The aircraft 100 shown in FIG. 1C may have the same or similar features and/or functionalities as the embodiment of the aircraft 100 shown in FIGS. 1A-1B, and vice versa. In particular, the rotor assemblies 230 may be the same or similar as the rotor assemblies as described with respect to FIGS. 1 A-1B.

[0106] As shown in FIG. 1C, the wingtip rotor assemblies 230 are oriented in a direction to provide thrust for forward flight. The rotor assemblies 230 may be rotated to provide a counter torque during takeoff and landing, as described with respect to FIGS. 1A- 1B. The propulsion system 105 may be a prime power system, an electric power system, or as shown a hybrid system that includes two electrical power subsystems 200 operatively connected to the prime power subsystem 300 and to respective main rotors 210. The main rotors 210 may each include the respective blades 212, 214, 216, which may or may not have features for changing geometric shape. The prime power subsystem 300 is operatively connected to the electrical power subsystems 200 to provide a hybrid propulsion power source to controllably rotate the main rotors 210. Each main rotor 210 provides vertical lift to the aircraft 100 during the vertical flight phases (e.g. takeoff and landing) and provides horizontal thrust to the aircraft 100 during the horizontal flight phase, as described above with respect to the embodiments of FIGS. 1A-1B having one main rotor 210.

[0107] The center wing 110 includes a left portion 140 and right portion 150. An inner end of the left portion 140 is connected to a left side of a body 301 of the prime power subsystem 300, and the opposite, outer end of the left portion 140 is connected to a right side of the electrical power subsystem 200 that is on the left side of the aircraft 100. The left wing 120 is connected to the left side of the electrical power subsystem 200 that is on the left side of the aircraft 100. An inner end of the right portion 150 is connected to a right side of the body 301 of the prime power subsystem 300, and the opposite outer end of the right portion 150 is connected to a left side of the electrical power subsystem 200 that is on the right side of the aircraft 100. The right wing 130 is connected to the right side of the electrical power subsystem 200 that is on the right side of the aircraft 100.

[0108] The electrical power subsystems 200 are located in housings 160 and 170. The electrical power subsystem 200 on the left is housed inside housing 160, and the electrical power subsystem 200 on the right is housed inside housing 170. The housings 160, 170 provide structural protection and containment for various components of the respective electrical power subsystems 200. The housings 160, 170 may be formed of composite, aluminum, other metals, other suitable materials, or combinations thereof. The housings 160, 170 connect the various portions of the wings, as described above. The housings 160, 170 are each, respectively, connected to the left and right tails 180, 190.

[0109] FIGS. 1D-1E are schematics of embodiments of hybrid propulsion systems 401 A and 401B respectively, having electric and prime power aspects, that may be used with the tail-sitter aircraft of FIGS. 1A-1B. As shown in FIGS. 1D-E, the main rotor 210 may be operatively connected to a gear box 403 which is connected to the generator 420. One or more engines 402, as shown there may be two engines 402, are operatively connected to the gear box to provide actuation. There may be one, three, or more engines 402. The generator 420 may receive and in some cases store (e.g. with batteries) electrical power and provide the electrical power. The systems 401 A, 401B include the wingtip rotor assemblies 230 each having the motor 233 and propeller 240. As shown in FIG. 1D, the generator 420 may provide power to the motors 233 of the wingtip rotor assemblies 230 to rotate the propellers 240. The power from the generator may also rotate the wingtip rotor assemblies 230 about their respective lateral axes. As shown in FIG. 1E, the system 401B may not have the motors 233 connected to the generator 420. In such case, the motors 233 may be powered by separate power subsystems, use batteries, etc.

[0110] As noted in FIGS. 1D and 1E, various options for using the systems 401 A, 401B may be implemented. The systems 401 A, 401B may have one engine 402 off during loiter or cruise phases of flight and on during takeoff and landing phases. The other engine 402 may always be on and spin the generator and rotor 210 during loiter and cruise phases. These are just examples and other approaches may be employed. [0111] The propulsion systems 401 A, 401B provide VTOL power required while also providing significant mission flexibility by providing a large amount of electrical power available on demand. The propulsion system architectures described in FIGS. 1D-1E provide significant redundancy while efficiently distributing thrust production where it is most beneficial. The propulsion system configurations can easily provide increased power to payloads since the generator and associated power electronics may be increased in size as needed. In addition, the power used by the main rotor 210 to generate thrust during cruise can be offloaded to the wingtip rotors 240 to provide significantly higher power with increased aerodynamic efficiency.

[0112] Several advantages exist for a two-engine 402 configuration. For example, to minimize penalties associated with VTOL vehicles where very low partial power is required during loiter, one engine 402 may be turned off in forward flight and the vehicle is ideally operated at altitudes over 15,000 ft. The two engines 402 make it unlikely that the main rotor 210 will be unpowered at any time since upon engine 402 failure a restart of either engine 402 would enable cruise flight to be maintained. In addition to attempts made to air start the engines 402, the main rotor 210 is capable of an autorotation-assisted landing if both engines 402 were to stop. Although air-restarts would be attempted first, the rotor 210 inertia and an auxiliary power unit (APU) (e.g., a 30-50kW APU) may help provide generator power to restart the engine 402 in case of an engine failure. In some embodiments, a 2-3% exhaust jet thrust available from each top-wing mounted engine may be used to help limit pitch and yaw trim loads imposed on the rotor 210.

[0113] In some embodiments, the propulsion systems 401 A, 401 B shown in FIGS. 1D-1E may be used with an aircraft having two or more main rotors 210, such as the aircraft 100 of FIG. 1C having three rotors 210. For example, there may be three rotors and gearboxes mechanically connected to respective engines 402, etc. Thus the configurations of FIG. 1D- 1E may be modified as needed for aircraft 100 having two or more main rotors 210.

[0114] FIG. 1F is a table showing an embodiment of control techniques for the various phases of flight that may be used with the aircraft 100 of FIGS. 1A-1B, or other aircraft. The flight controls for both VTOL and cruise are shown in FIG. 1F. For example, during VTOL, the aircraft 100 may use traditional helicopter rotor controls, providing full cyclic and collective authority with a swashplate. This may ensure sufficient control authority is possible to pitch the aircraft 100 during transition and conversion when tail-sitters can have difficulty due to significant flapping of the rotor blades. About ten degrees of clearance may be provided to accommodate this flapping. In some embodiments, more than four, more than six, more than eight, more than ten, more than twelve or more degrees of clearance may be provided. The full rotor authority also provides effective low speed control authority in VTOL, transition and conversion. Asymmetric deflections of the inboard ailerons outboard of the nacelle provide the ability to trim with the downwash induced by the rotor during hover, reducing the thrust required from the motors (e.g., electric motors) to trim the rotor torque. During fixed wing flight, trim is provided by the rotor. This enables significant flexibility in center of gravity (CG) location if needed for mission flexibility. The rotors 240 will also provide responsive control authority to augment the main rotor 210 cyclic if needed.

[0115] In some embodiments, the wingtip rotor assemblies 230 may be pointing forward in forward flight to contribute to forward thrust forces. The aircraft 100 may travel at speeds of 100 or more, 150 or more, 200 or more, 250 or more, 300 or more, 350 or more, 400 or more, 450 or more, 500 or more miles per hour. The rotor assemblies 230 may provide for enhanced aerodynamic characteristics of the aircraft 100 at these and other speeds. For example, in forward flight, the rotor assemblies 230 may create up wash on the wingtip or downstream air flow from the rotors 240 that reduces drag due to lift. Drag due to lift may be caused by higher pressure air under the wing 119 spilling laterally over the wingtip into the lower pressure region above the wingtip. This causes a downward pressure and induces a wingtip vortex creating a drag force on the airplane. The upwash from the rotor assemblies 230 may act as a buffer or separation that prevents or reduces the higher pressure air under the wing from spilling over into the lower pressure region over the wing, thus reducing the drag due to lift. The rotor assemblies 230 may also act like end plates, which are structures extending upward and/or downward from the wingtips that further prevent such spilling of air.

[0116] FIG. 1G is a table showing an embodiment of aspects of three different schemes for rotating the wingtip rotor assemblies 230. The schemes 1, 2 and 3 of FIG. 1G may correspond to the configurations of the assemblies 230- 1 , 230-2 and 230-3 shown in FIGS. 1H, II and 1 J, respectively. The assemblies 230 may include an actuator 237, a motor mount 235, the motor 233 and the rotor 240. The actuator 237 may be a linear actuator with pin-pin connect. The actuator 237 pivots or rotates the motor 233 about the lateral axis to thereby reorient the thrust vector. Each assembly 230 may include a single actuator 237. In some embodiments, there may be multiple actuators 237. In some embodiments, the motor 233 and a rotatable portion 236, such as a nose cone, may be rotated. Thus a fixed portion 234, such as a nacelle, may not be rotated. This may provide a structurally efficient configuration with enhanced control of the thrust angle. The assemblies 230 may be all electric as described herein, for example with an electrical connection extending to the wingtip from the generator or other electrical power source. In some embodiments, other types of rotation mechanisms besides electric actuators may be used, such as hydraulic which may save weight. Further, the rotating portions of the assemblies 230 in some embodiments may rotate in both directions about the lateral axis, for example rotate in a first direction to provide a thrust vector component in the +Z direction and then rotate in a second direction to provide a thrust vector component in the -Z direction.

[0117] As shown in FIG. 1H and described in Scheme 1 of FIG. 1G, the assembly 230 may include the actuator 237 aligned with the thrust generated by the rotor 240. The pivot interface with the nacelle may be at or near the maximum height of the electric motor’s thrust line. As shown in FIG. II and described in Scheme 2 of FIG. 1G, the assembly 230 may include the actuator aligned with the efficient structure. The pivot or rotation of the assembly 230 may be aligned with the efficient structure. As shown in FIG. 1 J and described in Scheme 3 of FIG. 1G, the assembly 230 may include the actuator positioned to provide the maximum thrust height change. The pivot interface with the nacelle may be as shown and provide minimum downwash on the wing 119.

[0118] FIGS. 1K-1F are sequential views of a wingtip rotor assembly shown respectively in a takeoff orientation and a cruise orientation. The assembly 230 may include the body 232, such as a nacelle as shown, at the wingtip region 122 of the wing 119. The thrust vector TV generated by the assembly 230 may be in either direction as shown. For example the rotor 240 (shown schematically in outer swept diameter of the blades) may be a pusher or a puller type rotor 240. The body 232 may rotate near but behind the center of mass of the body 232, as shown. The actuator for rotating the assembly 230 about the lateral axis may be in the wingtip of the wing 119 or in the body 232.

[0119] FIG. 1M is a partial schematic of the aircraft 100 showing example thrust vector components and resulting torques that may be created by the various aircraft and systems described herein. The aircraft 100 body is removed for clarity. As shown, schematics of the main rotor 210 and wingtip rotor assemblies 230 A and 230B are indicated. The XYZ reference system is shown. The main rotor 210 may rotate to create a thrust vector (TV) in the direction shown, which may be aligned with the X axis in this example. The torque this rotation creates is indicated in the figure and acts about the X axis in the clockwise direction as viewed in the figure.

[0120] The first rotor assembly 230A is rotated about a lateral axis (e.g., the Y axis or an axis parallel thereto) to the orientation shown, and in this orientation rotates the rotor to create a first thrust vector TV1 in the direction indicated. The vector TV1 has components Zl and XI that are parallel respectively to the Z and X axes of the reference XYZ system. In this example, the Y component of TV1 is zero. The second rotor assembly 230B is rotated about a lateral axis to the orientation shown, and in this orientation rotates the rotor to create a second thrust vector TV2 in the direction indicated. The vector TV2 has components Z2 and X2 that are parallel respectively to the Z and X axes of the reference XYZ system. In this example, the Y component of TV1 is zero.

[0121] As shown, the XI and X2 components are in the same direction and contribute to upward thrust of the aircraft. The thrust vector TV from the main rotor is also in the positive X direction and contributes to the upward thrust of the aircraft. Further, the components Zl and Z2 are in opposite directions relative to each other. The component Zl is in a positive Z direction and the component Z2 is in a negative Z direction, relative to the XYZ reference system as oriented. The Zl and Z2 components thus contribute to a second torque about the X axis of the XYZ reference system. This second torque is shown as a“counter torque” in the figure, because it is in an opposite direction the torque create by the spinning main rotor 210. Thus the Z components of the rotor assemblies 230A, 230B create a torque that counters the torque from the main rotor 210. The counter torque may also be equal in magnitude to the torque created by the main rotor 210. Thus as the aircraft 100 takes off, it may maintain a constant or approximately constant rotational orientation relative to the X axis. This may allow for controlled takeoff, for stability, for taking off in tight or small locations that require the aircraft to maintain its angular orientation to avoid collisions, etc.

[0122] In some embodiments, the rotor assemblies 230A, 230B may be oriented such that no upward thrust is provided as shown. For example, the assemblies 230A, 230B may have the orientations shown in FIG. 1B, such that the X and Y components of the thrust vectors are zero, and only a Z component is created by the respective thrust vectors TV 1 , TV2. Further, the torque created by the main rotor 210 may be the result of multiple main rotors 210 such that the torque is a net torque. The rotor assemblies 230A, 23 OB may counter this net torque as described.

[0123] FIG. 2 is top perspective view of the tail-sitter aircraft 100 of FIGS. 1A-1B with a hybrid propulsion system 105 and a single main rotor 210, with portions of the body of the aircraft 100 shown transparently for clarity. Thus, portions of the electrical power subsystems 200 and prime power subsystem 300 are visible. The electrical and prime power subsystems 200, 300 may be used controllably in varying amounts depending on the phase of flight, such as takeoff, horizontal flight, landing, or maneuvers.

[0124] The tail-sitter aircraft 100 has a fuselage that includes the body 301, a wing 198, and a hybrid propulsion system 105. The fuselage may also include the housings 160, 170 and the tails 180, 190. The fuselage has a nose end, that includes the forward portions of the components of the fuselage, and a tail end that includes the rearward portions of the components of the fuselage. In some embodiments, there may only be a single, central structure, such as the body 301, where the fuselage only includes the body 301. In some embodiments, there may be more than one body 301 and/or more than two housings 160, 170, and the fuselage would include all of these components with the nose end being the forward portions thereof and the tail end being the rearward portions thereof. The tail end and/or other features may support the aircraft 100 on the ground with the nose end oriented vertically or generally vertically. The hybrid propulsion system 105 includes the electrical power subsystem 200 coupled with the main rotor 210, an electrical energy store 211 coupled with the electrical power subsystem 200, and the prime power subsystem 300 coupled with the electrical power subsystem 200.

[0125] The wing 198 is coupled with the body 301. The wing 198 provides lift during the horizontal flight phase. The wing 198 extends between the body 301, which houses the prime power subsystem 300, and the housings 160, 170, which house the electrical power subsystems 200. The wing 198 extends beyond the electrical power subsystems 200. In some embodiments, the wing 198 may extend to electrical power subsystem 200 and not beyond. The wing 198 may hold fuel. The wing 198 may have one or more fuel tanks for holding fuel. The fuel type depends on the type of engine used in the prime power subsystem 300. The shape and length of the wing 198 illustrated in FIG. 2 is merely exemplary and the wing 198 can have any suitable shape and length for flight. The wing 198 can include the left wing 120, the left portion 140, the right portion 150, and the right wing 130 described with respect to FIG. 1A.

[0126] Typically, the sweep of the wing 198 is about zero degrees about the quarter chord of the wing 198 for peak aerodynamic efficiency during subsonic flight and peak structural efficiency. For a tail-sitter aircraft 100, the wing sweep may have a deviation of about +20 degrees from the baseline of zero sweep to prevent wing tip strike during takeoff and landing, to distance the outboard wing leading edge away from the downward flow induced by the blade assembly, and/or to correct the center of gravity and aerodynamic center placement. As illustrated in the embodiment in FIG. 2, the wing 198 is swept forward. In some embodiments, the wing 198 may be swept differently (e.g. straight or backward).

[0127] As illustrated in the embodiment in FIG. 2, the electrical power subsystems 200 are located inside the left and right housings 160 and 170. The positions of the electrical power subsystems 200 illustrated in FIG. 2 are merely exemplary, and the electrical power subsystems 200 could have any suitable position inside or outside the housing 160, 170 based on spacing, power, or safety concerns.

[0128] The electrical power subsystem 200 includes an electric power generation system 202. The electric power generation system 202 includes an electric motor 220 operatively connected to a generator 320 as described herein, for example as shown in and discussed with respect to FIG. 3. The electric power generation system 202 is coupled with the main rotor 210. The electric power generation system 202 supplies an increased electrical power to rotate the main rotor 210 at a first speed during the vertical takeoff and landing phases and a reduced electrical power to rotate the main rotor 210 at a second speed during the horizontal flight phase. The first speed is greater than the second speed. In some embodiments, the speed during vertical takeoff is about the same as the speed during vertical landing. In some embodiments, the speed during vertical takeoff is greater or lower than the speed during vertical landing. The speed is measured as the revolutions per minute of the blade assembly. The electric power generation system 202 supplies an increased electrical power during the vertical phases than during the horizontal flight phase. [0129] The aircraft 100 includes an energy store 211, which may be, for example, one or more battery packs. The energy store 211 can have one or more electric batteries. An electric battery can have one or more electrochemical cells, for example, alkaline, lead-acid, lithium-ion, nickel-cadmium, nickel-zinc, nickel metal hydride, zinc-carbon, or the like. The battery can be single use or rechargeable.

[0130] The energy store 211 is coupled with the electric power generation system 202. The energy store 211 provides electrical energy to the electric power generation system 202 during the vertical takeoff and landing phases, which require an increased electric power output compared to the horizontal flight phase. The energy store 211 stores electric energy produced by the electric power generation system 202 during the horizontal flight phase, which requires a reduced electric power output compared to the vertical phases. In some embodiments, the energy store 211 is considered part of the electrical power subsystem 200. In some embodiments, a generator, such as the generator 320 described herein, is considered part of the electrical power subsystem 200.

[0131] The prime power subsystem 300 may be located entirely or partially inside the body 301, for example with a“series hybrid configuration,” described herein. In other configurations, such as a“parallel hybrid configuration” described herein, the prime power subsystem 300 may be co-located with the electric power subsystem 200. The position of the prime power subsystem 300 is merely exemplary; the prime power subsystem can have any suitable position inside or outside the body 301 based on spacing, power, or safety concerns. The prime power subsystem 300 includes a prime power generation system 302. Details for the prime power subsystem 300 are shown in and discussed in greater detail herein, for example with respect to FIG. 3.

[0132] FIG. 3 is a bottom view of the prime power subsystem 300 and the electric power generation system 202 with the body 301 shown transparently for clarity. The prime power subsystem 300 includes a prime power generation system 302. The prime power generation system 302 may be an engine 310, such as a combustion engine. The engine 310 can use various fuels, for example, fossil fuels, natural gas, coal, petroleum, gasoline, diesel, fuel oil, renewable fuels, biofuels, biodiesel, bioethanol, methanol, hydrogen, or the like. The engine 310 may be a diesel engine. The engine 310 may be a jet engine, turbo-fan engine, turbo-prop engine, piston engine, spark ignition, compression ignition, fuel cell, or the like. In some embodiments, an efficient loiter power source is selected, such as a fuel cell or similar source where the prime power subsystem 300 is very efficient for loiter which may be incapable of providing the peak power required by VTOL, but could provide electrical power which could be leveraged by VTOL.

[0133] The prime power generation system 302 includes a generator 320. The generator 320 is operatively coupled with the engine 310 and converts the mechanical energy created by the engine 310 to electrical energy. The generator 320 may be an alternator, direct current generator, alternating current generator, induction generator, linear electric generator, variable speed constant frequency generator, or the like. In some embodiments, an alternating current (AC) motor may be used as the generator 320. In some embodiments, the generator 320 may instead or in addition be considered part of the combustion engine 310, for example, as a single unit engine-generator. In some embodiments, the generator 320 may instead or in addition be considered part of the electrical power subsystem 200, such that the electric power generation system 202 may comprise the generator 320 and the electric motor 220. Thus, the description of the generator with respect to either the electric or prime power subsystems is not limiting in that regard.

[0134] The prime power subsystem 300 may include one or more sensors 330. The sensor 330 is a sense-and-avoid assembly. The sensor 330 is able to detect and avoid intruding aircraft or objects at least within a 3 mile radius with a field of regard of 270 x 30 degrees. The sensor 330 is used for autonomous flight guidance, navigation, and control. In the embodiment illustrated in FIG. 3, the sensor 330 is positioned near the front of the body 301. This position is merely exemplary and the sensor 330 may be housed in other locations of the aircraft 100, for example, in the wing 198 or the housing 160 or 170. The sensor 330 may be positioned in the front, rear, top, or bottom half of the aircraft 100. The sensor 330 may have its own power source. The sensor 330 may be powered by other power sources, such as the electric power generation system 202 or energy store 211.

[0135] The prime power subsystem 300 may include one or more instruments 340, such as an optical instrument. In the embodiment illustrated in FIG. 3, the instrument 340 is shown as a laser designator, but the instrument 340 may be another type of instrument, for example, visual camera, infrared sensor, ultraviolet light sensor, night vision device, laser light, heat detector, photometer, refractometer, reflectometer, or the like. In the embodiment illustrated in FIG. 3, the instrument 340 is positioned near the rear of the body 301. This position is merely exemplary and the optical instrument 340 may be housed in other locations of the aircraft 100, for example, in the wing 198 or the housing 160 or 170. The instrument may be positioned in the front, rear, top, or bottom half of the aircraft 100.

[0136] FIG. 4 is a schematic of an embodiment of a hybrid propulsion system configurations, having electric and prime power subsystems, that may be used with the aircraft 100. The hybrid propulsion system may be used as the hybrid propulsion system 105 described herein. Further, although details of the various components of each configuration may be described with respect to one or another of the configurations shown in FIGS. 4, it is understood that the components described with respect to a particular configuration may have the same or similar features as the components of the other configurations, and vice versa.

[0137] FIG. 4 is a schematic of an embodiment of a hybrid propulsion system 400A that may be used with various aircraft having wingtip rotor assemblies 230 and multiple main rotors 210, such as the aircraft 100 of FIG. 1C. Modifications to the system may be made to use with the aircraft of FIG. 1C, for example with an extra electric motor and gear box, as further described, or of FIGS. 1 A-1B for example with a single electric motor and gear box, as further described.

[0138] The hybrid propulsion system 400A may have the same or similar features and/or functionalities as the hybrid propulsion systems 105, 400, and vice versa. The hybrid propulsion system 400 A includes a prime power subsystem 300A and an electrical power subsystem 200 A, as indicated. The prime power subsystem 300 A and electrical power subsystem 200A may have the same or similar features and/or functionalities as, respectively, the prime power subsystem 300 and electrical power subsystem 200, and vice versa. In general, as used herein, and unless otherwise indicated explicitly or by context, callouts including a suffix such as“’” (apostrophe),“A”,“B” etc. are understood that they may have the same or similar features and/or functionalities as similar callouts having different or no suffixes, such as 300 and 300A, 200 and 200A, etc.

[0139] The prime power subsystem 300 A includes an engine 410A, a gear box 415 A, and a generator 420A. In FIG. 4, the engine 410A is shown as a diesel engine, but the engine 410A may be any type of engine, for example, internal combustion engine, a jet engine, turbo-fan engine, turbo-prop engine, piston engine, spark ignition, compression ignition, fuel cell, or the like. The gearbox 415A operatively connects the engine 410A to the generator 420A. The gearbox 415A transfers the power from the engine 410A to the generator 420A. In some embodiments, the gearbox 415A may be in combination with the engine 410A and/or generator 420A. In some embodiments, the transfer of power may be done without a gearbox, such as a diesel-electric transmission or gas-electric transmission.

[0140] The electrical power subsystem includes an energy store 430 A, electric motors 440 A, 440A’, gearboxes 450A, 450A’, attached to blade assemblies 460A, 460A’ (may also be referred to as propeller or rotor), and may include the generator 420A. In some embodiments, there may be only one or more than two of the electric motor, gear box, and blade assembly. The gearboxes 450A, 450A’ may adjust the outputs from the electric motors 440 A, 440A’ respectively to adjust the rotational speed for the blade assemblies 460A, 460 A’ respectively, where the speed may be measured in rotations or revolutions per minute (RPM). The energy store 430A, which may be one or more batteries, is coupled to both electric motors 440 A, 440A’. Each gearbox 450A, 450A’ is coupled to its own blade assembly 460A, 460A’, respectively. The blade assemblies 460A, 460A’ may be analogous to the main rotor 210.

[0141] The electric power subsystem 200A includes the generator 420A coupled with the gear box 415A of the prime power subsystem 300A, and the generator 420A is also coupled with the electric motors 440 A, 440 A’ and the energy store 430A. The electric motors 440A, 440A’ are coupled with respective blade assemblies 460A, 460A’ via the respective gear boxes 450A, 450A’.

[0142] The prime power subsystem 300A provides prime power to the generator 420A for the production of electrical power. The generator 420A supplies the electrical power to the electric motors 440A, 440A’ to rotate the blade assemblies 460 A, 460A’ . The generator 420A supplies increased electrical power to the electric motors 440 A, 440A’ to rotate the blade assemblies 460A, 460A’ at a first speed during the vertical takeoff and/or vertical landing phases. The electrical power provided by the generator 420A during vertical phases of flight is greater than the electrical power provided by the generator 420A during the horizontal or loiter phases of flight.

[0143] The generator 420A is configured to supply the increased electrical power to the electric motors 440 A, 440A’ to rotate the blade assemblies 460 A, 460A’ at high speed during the vertical takeoff and vertical landing phases. “High” speed is the speed relative to horizontal flight and is a speed that, along with a corresponding lower pitch (relative to horizontal flight), allows the aircraft to lift off the ground. Pitch here refers to the pitch of a section or sections of the blade, which may be changed by twist of the blade and/or by rotation of the blade to uniformly change the pitch. A higher pitch of the blade assemblies 460 A, 460A’ can generate greater propulsive forces but only until portions of the blade assemblies 460A, 460A’ stall, and it thus generates that increased propulsive force with decreased overall energy efficiency since more drag is produced and therefore more power is required to rotate the blade assemblies 460A, 460A’ at a sufficiently fast speed. The greater pitch is therefore used to account for local angle of attack changes when the blade assemblies 460 A, 460A’ experience increased flow velocity parallel to the axis of rotation of the blade assemblies 460A, 460A’. The blade assemblies 460 A, 460A’ can thus more efficiently generate thrust with greater pitch at higher forward speeds of the aircraft 100, such as during horizontal flight, and the lower pitch is thus used during vertical takeoff and landing when the forward movement of the aircraft 100 is slower compared to horizontal flight.

[0144] The electrical energy store 430 A is coupled with the electric motors 440A, 440A’. The electrical energy store 430A provides the increased electrical energy to the electric motors 440A, 440A’ during the vertical takeoff and landing phases. The blade assemblies 460 A, 460A’ rotate at a higher speed during vertical takeoff and landing phases, as described, so the electric motors 440A, 440A’ require an increased electrical energy compared to the electrical energy needed during horizontal flight. The electrical energy store 430A is coupled with the generator 420 A. The electrical energy store 430A stores electrical energy produced by the generator 420A during horizontal flight. Various other embodiments of a hybrid propulsion system may be used for the various aircraft 100 described herein, including for example, those described in U.S. patent application no. 15/792,490, filed October 24, 2017, and titled“Tail-Sitter Aircraft with Hybrid Propulsion.” The system 400A may be modified for aircraft having a single main rotor 210 or three or more main rotors 210. There may only be a single electric motor 44A connected to the generator 420A and batteries 430A. Or there may be three or more electric motors 440A, 440A’ etc. connected to the generator 420A and batteries 430A. There may be correspondingly a single gear box 450A and rotor 460 A, or there may be multiple gear boxes, 450A, 450A’, etc. and rotors 460A, 460A’ , etc. [0145] FIG. 5 is a flowchart showing an embodiment of a method 500 for operating a hybrid propulsion system of a tail-sitter aircraft for vertical takeoff, horizontal flight and vertical landing. The method 500 may be used with various tail sitter aircraft 100 having the hybrid propulsion system 105 or 400A and wingtip rotor assemblies 230. The method 500, or portions thereof, may be employed with other methods described herein, for example the methods 1710 and 1730 shown and described with respect to FIGS. 14A-14B.

[0146] The method 500 begins with step 510 wherein the aircraft is positioned with the nose up for vertical takeoff. Step 510 may include, for example, positioning the aircraft 100 on its tails 180, 190 and/or on other structures of the aircraft 100. In step 510, the aircraft 100 may be positioned in a variety of suitable ways, including landing on the tail from a prior flight, positioned by humans and/or machines, etc. The nose may be pointing vertically or off- vertically.

[0147] The method 500 next moves to step 520 wherein a high (i.e. peak) power is supplied for a short duration from the hybrid propulsion system 105. Step 520 may include, for example, providing high power from the hybrid propulsion system 105 for takeoff. The power provided from the hybrid propulsion system 105 in step 520 for takeoff is expected to be higher than the power provided from the hybrid propulsion system 105 for horizontal flight. The duration that power is provided for takeoff in step 520 is expected to be shorter than the duration that power is provided for horizontal flight. In some embodiments of step 520, the hybrid propulsion system 105 may provide the high power using the prime power subsystem 300 A, 300B, 300C, 300D or 300E, the electrical power subsystem 200A, 200B, 200C, 200D or 200E, or a combination of the prime and electrical power subsystems. In step 520, the prime power subsystem, such as the prime power subsystem 300 A, 300B, 300C, 300D or 300E, and the electrical power subsystem, such as the electrical power subsystem 200A, 200B, 200C, 200D or 200E, may provide peak power. The high power provided is sufficient to rotate the main rotor 210 at a high enough rate to lift the aircraft 100 from the ground. The duration is sufficient to lift the aircraft 100 to a desired height.

[0148] The method 500 next moves to step 530 wherein the aircraft takes off vertically. “Vertically” in step 530 includes the vertical flight phase situations described herein, such as when the blade thrust provides the primary lifting force, etc. Step 530 may include, for example, the aircraft 100 raising to a height from its position on its tails 180, 190. As the aircraft 100 takes off, the aircraft generally stays in a vertical alignment, where the front of the aircraft 100 is generally pointed away from the ground. In some embodiments, the longitudinal axis of the aircraft 100, shown in FIG. 1 A, stays generally vertical during takeoff. Step 530 includes the blade assembly, such as the main rotor 210, rotating the blades thereof to produce the propulsive lifting forces.

[0149] The method 500 then moves to step 540 wherein the aircraft is rotated for horizontal flight. “Horizontal” in step 540 includes the horizontal flight phase situations described herein, such as when the wings provide the primary lifting force, etc. Step 540 may include, for example, rotating the aircraft 100 from its vertical takeoff alignment to a horizontal alignment for horizontal, sustained flight. In some embodiments, the longitudinal axis of the aircraft 100, shown in FIG. 1A, stays generally perpendicular to vertical during horizontal flight. The aircraft may be rotated in step 540 by cyclically altering the twist and/or pitch of the main rotor 210, such as with helicopter rotor flight controls, to cause a forward moment or torque on the aircraft 100 that causes it to rotate. In some embodiments, the aircraft may be rotated in step 540 by rotation of the blade assemblies, such as rotation of the hub 218 of the main rotor 210. Further details of how a blade may twist and thus change shape to effect rotation of the aircraft 100 in step 540 are described in further detail, for example, in U.S. patent application no. 15/792,542, filed October 24, 2017, and titled“Shape Changing Aircraft Blade,” the entire content of which is incorporated herein by reference in its entirety for all purposes and forms a part of this specification. The aircraft may be rotated in step 540 by using shape-changing blades and/or control surfaces of the aircraft 100. Further details of how shape-changing blades and/or control surfaces of the aircraft 100 may be used in step 540 to effect rotation of the aircraft are described in further detail herein, for example with respect to FIG. 10B.

[0150] The method 500 then moves to step 550 wherein a low power is supplied for a long duration from the hybrid propulsion system 105. Step 550 may include, for example, providing low power from the hybrid propulsion system 105 for horizontal, sustained flight. The power provided from the hybrid propulsion system 105 for horizontal flight is expected to be lower than the power provided from the hybrid propulsion system 105 for takeoff. The duration that power is provided for horizontal flight is expected to be longer than the duration that power is provided for takeoff. In some embodiments, the hybrid propulsion system 105 may provide the low power using the prime power subsystem 300 A, 300B, 300C, 300D or 300E, the electrical power subsystem 200 A, 200B, 200C, 200D or 200E, or a combination of the prime and electrical power subsystems. The low power provided is sufficient to keep the aircraft 100 aloft. The duration is sufficient for the aircraft 100 to go a desired distance or desired amount of time.

[0151] The method 500 then moves to step 560 wherein the aircraft flies horizontal. “Horizontal” in step 560 includes the horizontal flight phase situations described herein, such as when the wings provide the primary lifting force, etc. Step 560 may include, for example, the aircraft 100 flying in a generally level alignment. In some embodiments, the longitudinal axis of the aircraft 100, shown in FIG. 1A, stays generally perpendicular to vertical during horizontal flight.

[0152] The method 500 then moves to step 570 wherein the aircraft is rotated for vertical landing. “Vertical” in step 570 includes the vertical flight phase situations described herein, such as when the blade thrust provides the primary lifting force, etc. Step 570 may include, for example, rotating the aircraft 100 from its horizontal flying alignment to a vertical alignment for vertical landing. In some embodiments, the longitudinal axis of the aircraft 100, shown in FIG. 1A, is rotated so that the front of the aircraft is pointed up and the rear of the aircraft is pointed towards the ground so that the aircraft 100 can land on its tail. The aircraft may be rotated in step 570 by cyclically altering the twist and/or pitch of the main rotor 210, such as with helicopter rotor flight controls, to cause a rearward moment or torque on the aircraft 100 that causes it to rotate. In some embodiments, the aircraft may be rotated in step 570 by rotation of the blade assemblies, such as rotation of the hub 218 of the main rotor 210. Further details of how a blade may change shapes to effect rotation of the aircraft in step 570 are described in further detail, for example, in U.S. patent application no. 15/792,542, filed October 24, 2017, and titled“Shape Changing Aircraft Blade,” the entire content of which is incorporated herein by reference in its entirety for all purposes and forms a part of this specification. The aircraft may be rotated in step 570 by using shape-changing blades and/or control surfaces of the aircraft 100. Further details of how shape-changing blades and/or control surfaces of the aircraft 100 may be used to effect rotation of the aircraft in step 570 are described in further detail herein, for example with respect to FIG. 10B. [0153] The method 500 then moves to step 580 wherein a high power is supplied for a short duration from the hybrid propulsion system 105. Step 580 may include, for example, providing peak or high power, as described herein, from the hybrid propulsion system 105 for landing. The power provided from the hybrid propulsion system 105 in step 580 for landing is expected to be higher than the power provided from the hybrid propulsion system 105 for horizontal flight. The duration that power is provided for landing in step 580 is expected to be shorter than the duration that power is provided for horizontal flight. In some embodiments, in step 580 the hybrid propulsion system 105 may provide the high power using the prime power subsystem 300 A, 300B, 300C, 300D or 300E, the electrical power subsystem 200A, 200B, 200C, 200D or 200E, or a combination of the prime and electrical power subsystems. The high power provided in step 580 is sufficient to lower the aircraft 100 to the ground. The duration is sufficient for the aircraft 100 to land. The power provided and duration for landing in step 580 may be the same or substantially similar to the power provided and duration for takeoff, such as in steps 520 and 530. In some embodiments, the power provided and/or duration for landing in step 580 may be less than for takeoff. In some embodiments, the power provided and/or duration for landing in step 580 may be greater than for takeoff. In some embodiments, the power provided may be of a duration sufficient for an autorotation landing only.

[0154] The method 500 then moves to step 590 wherein the aircraft lands vertically with the nose up. Step 590 may include, for example, the tails 160, 170 of aircraft 100 touching the ground as the aircraft 100 lands in a vertical alignment. Various other configurations of the tail may be used, such as those discussed with respect to FIGS. 10A-10C. The aircraft 100 may land in step 590 without landing on the tail, for example other landing structures. The nose in step 590 may be pointing vertically up or off-vertical.

[0155] FIG. 6A is a flowchart showing an embodiment of a method 600 for operating the hybrid propulsion system 105 of the tail-sitter aircraft 100 having wingtip rotor assemblies 230 that may be used for the vertical takeoff and landing phases of flight. The method 600 may thus be used for some of the steps of the method 500, for example the steps 520 and/or 580. The method 600, or portions thereof, may be employed with other methods described herein, for example the methods 1710 and 1730 shown and described with respect to FIGS. 14A-14B. [0156] The method 600 may begin with step 610 wherein a high prime power is supplied. In some embodiments, the prime power subsystem 300 A, 300B, 300C, 300D or 300E supplies a first prime power to the aircraft engine during takeoff. It is expected that the first prime power during takeoff is greater than a second prime power provided during horizontal flight because it is expected that more power is needed to generate the sufficient propulsive lifting force for takeoff.

[0157] The method 600 then moves to step 620 wherein a high electric power is supplied from the electric energy store. In some embodiments, the high electric power is supplied from the electrical power subsystem. In some embodiments, an electric power source (e.g. electrical power subsystem 200A, 200B, 200C, 200D or 200E and/or electric energy store 211) supplies a first electric power to the aircraft engine during takeoff. Steps 610 and 620 may be performed simultaneously to provide maximum power for takeoff or landing.

[0158] FIG. 6B is a flowchart showing an embodiment of a method 650 for operating the hybrid propulsion system 105 of the tail-sitter aircraft 100 having wingtip rotor assemblies 230 that may be used for the horizontal flight phase. The method 650 may thus be used for some of the steps of the method 500, for example, the step 550. The method 650, or portions thereof, may be employed with other methods described herein, for example the methods 1710 and 1730 shown and described with respect to FIGS. 14A-14B.

[0159] The method 650 may begin with step 660 wherein a low prime power is supplied. In some embodiments, the prime power subsystem 300 A, 300B, 300C, 300D or 300E supplies a second prime power to the aircraft engine during horizontal, sustained flight that is less than a first prime power provided for takeoff and landing.

[0160] The method 650 then moves to step 670 wherein the electric energy store is charged. In some embodiments, the generator 420A charges the energy store 430A (see FIG.

4)·

[0161] The method 650 then moves to step 680 wherein a low electric power is supplied. In some embodiments, the low electric power is supplied from the electrical power subsystem. In some embodiments, the electric power source (e.g. electrical power subsystem 200 A, 200B, 200C, 200D or 200E and/or electric energy store 211) supplies a second electric power to the aircraft engine during horizontal, sustained flight that is less than a first electric power provided for takeoff and landing. [0162] The hybrid propulsion system 105 can provide varying ratios of high to low power for takeoff and landing compared to horizontal flight. In some embodiments, the power provided for each of takeoff and landing may be multiples of the power provided for horizontal flight. For example, a first sum may be the sum of the first prime power and the first electric power provided during each of takeoff or landing. A second sum may be the sum of the second prime power and the second electric power provided during horizontal flight. The first sum is sufficient to provide vertical lift in an amount at least equal to a force due to gravity on the aircraft. The second sum is sufficient to sustain horizontal flight. The first sum is greater than the second sum. The first sum may be at least one-and-a-half, two, two-and-a-half, three, four, five, six, seven, eight or more times larger than the second sum. The first sum may be about 300 horsepower. The second sum may be about 60 horsepower. In some embodiments, the ratio of the first sum to the second sum is in the range of from about 1.5: 1 to about 8: 1. The first sum for takeoff and landing may use from about 150-450 horsepower. The second sum for horizontal flight may use from about 30-90 horsepower. These are merely some examples and other amounts, proportions, ratios, etc. may be used.

[0163] FIGS. 7A-7F are perspective views of an embodiment of a blade assembly 700 having shape-changing aircraft blades 720, 730 that may be used as the blades 212, 214, 216 for the main rotor 210 and/or the blades 242 of the wingtip rotor 240 of the aircraft 100. The aircraft 100 having wingtip rotor assemblies 230 may use shape-changing rotor blades for the main rotors 210 and/or wingtip rotors 240. The assembly 700 in FIGS. 7A-7C shows the blades 720,730 in a first configuration having a first pitch that may be used with the tail-sitter aircraft 100. The blade assembly 700 can have one or more than two blades. FIG. 7A illustrates an embodiment with the first blade 720 and the second blade 730, both attached to a hub 710. A“longitudinal directions” double-sided arrow is indicated, showing the dimension along which the term“longitudinal length” of the blade 720 refers, as used herein. The second blade 730 may have the same or substantially similar shape as the first blade 720. The second blade 730 may have a shape different from the first blade 720. The blades 720, 730 are connected to the hub 710 such that the blades 720, 730 can alter their pitch. The blade assembly 700 may have the same or similar features and/or functionalities as the other blade assemblies described herein, such as the main rotor 210, and vice versa. [0164] The blades 720, 730 have a twisted form to keep the thrust constant along the length of the blade because the velocity of a blade varies from the inner end near the hub 710 to the outer, opposite end. FIG. 7B depicts a top view of the outer portion of the blade 720 in a first configuration having a first pitch. FIG. 7C depicts a side view of the same outer portion of the blade 720 in the first configuration having the first pitch.

[0165] The blades 720, 730 may be pre-stressed to maintain the twist shape. The pre-stressing also facilitates changing the pitch due to piezo elements, as described herein. For example, the material of the blades 720, 730 may be biased to bend or twist in a certain direction. Application of a shape-changing force, such as a mechanical force induced by a current supplied to a piezo element, may twist the blade 720, 730 away from the direction in which it is biased due to pre-stressing. Removal of the shape-changing force, such as removal of a mechanical force due to removal of the current supplied to the piezo element, may cause the blade 720, 730 to twist back in the direction in which it is biased due to the pre-stressing. The materials may be pre-stressed in a number of suitable manners, such as bending, cyclic loading, progressively increased loading, progressively increased cyclic loading, orientation of materials (e.g. orientation of fibers within composite materials), etc.

[0166] The blade 720 of the aircraft has a first pitch for takeoff. The shape of the blade 720 may have the same or substantially similar pitch for landing. The shape, e.g. pitch, of the blade 720 for takeoff may be different from the shape, e.g. pitch, of the blade 720 for landing. The difference between the pitch angle for takeoff and the pitch angle for landing of the blade 720 may be about 2 degrees or less. The shape of the blade 720 is at a second pitch for horizontal flight. The second pitch is greater than the first pitch. The second pitch may be about 20 degrees to about 30 degrees greater than the first pitch. The propeller (also referred to as blade assembly or rotor) provides vertical lift to the aircraft during the vertical flight phases (e.g., takeoff and landing) and provides horizontal thrust to the aircraft during the horizontal flight phase. The blade 730 may have the same or similar features and/or functionalities as the blade 720.

[0167] The electrical and prime power subsystems, described herein, collectively rotate the blade assembly 700 at a relatively higher speed during the vertical takeoff and vertical landing phases and collectively rotate the blade assembly 700 at a relatively lower speed during the horizontal flight phase. The rotational speed of the blade assembly 700 during a vertical flight phase (e.g. takeoff and landing) is greater than the rotational speed of the blade assembly 700 during the horizontal flight phase. The rotational speed of the blade assembly 700 may be measured in revolutions per minute (RPM). The blade assembly 700 may function similar to a helicopter rotor during the vertical flight phases and as a non-tail-sitter aircraft propeller during the horizontal flight phase.

[0168] The change in shape of the blade assembly 700 may alter its pitch. The blade assembly 700 may be twisted to alter the pitch. The twist may be induced by piezoelectric material located within and/or on the blade assembly 700, as described herein. The blade assembly 700 may have internal composite plies. The composite plies may be any composite material, such as carbon fiber in epoxy, graphite fiber reinforced plastic, etc. The blade assembly 700 may further include a piezoelectric actuator adjacent the composite plies, as further described herein.

[0169] Changing the pitch alters the angle of attack of the aircraft blades 720, 730 and thus the vertical acceleration or climb rate of the vehicle. This control is also called collective as distinct from the cyclic control for lateral movement. The collective blade setting may be achieved through vertical movement of a swashplate. Cyclic and collective control are described in further detail, for example, in U.S. patent application no. 15/792,542, filed October 24, 2017, and titled“Shape Changing Aircraft Blade,” the entire content of which is incorporated herein by reference in its entirety for all purposes and forms a part of this specification.

[0170] The blades 720, 730 may be“feathered” to increase their angle of pitch by turning the blades to be parallel to the airflow. This may minimize drag from a stopped blade 720, 730 following an engine failure in flight.

[0171] FIGS. 7D-7F are perspective views of the shape-changing aircraft blades 720, 730 in a second configuration having a second pitch that may be used with the tail-sitter aircraft 100, for example for the horizontal flight phase. FIG. 7E is a top view of the outer end of the blade 720 in a second configuration having a second pitch. FIG. 7F is a side view of the outer end of the blade 720 in a second configuration having a second pitch. Thus, the blade 720 depicted in FIGS. 7B and 7C has a greater pitch than the blade 720 depicted in FIGS. 7E and 7F. [0172] FIG. 8 is a flowchart showing an embodiment of a method 900 for operating the shape-changing aircraft blades of the tail-sitter aircraft 100 in flight. While described with respect to particular aircraft and blades, the method 900 may be performed with the various aircraft and blades described herein, such as the aircraft 100 and blades 212, 214, 216, 242, 720, 720’, etc. The method 900, or portions thereof, may be employed with other methods described herein, for example the methods 1710 and 1730 shown and described with respect to FIGS. 14A-14B

[0173] The method 900 begins with step 905 wherein the aircraft 100 is positioned with the nose up for vertical takeoff. The step 905 may include, for example, positioning the aircraft 100 on its tails 180, 190 and/or on other structures for vertical takeoff. The aircraft 100 may be positioned in a variety of suitable ways, including landing on the tail from a prior flight, positioned by humans and/or machines, etc. In step 905 the nose may be pointing vertically or off-vertically.

[0174] The method 900 then moves to step 910 wherein a first electric current is supplied to a piezo element of a blade of the aircraft 100. Step 910 may include, for example, applying the first electric current from an energy store to the piezo element 740. The first electric current may induce a deformation of the piezo element 740 in step 910.

[0175] The method 900 then moves to step 915 wherein the shape of the blade is changed to a first shape having a first twist. Step 915 may include, for example, the piezo element 740 embedded in the blade 720 inducing a twist in the blade 720, thus altering the pitch of the blade 720. The twist may non-uniformly or uniformly change the pitch along the longitudinal or spanwise direction of the blade. In some embodiments, in step 915 the entire blade may be rotated, for example about the blades longitudinal axis, to uniformly change the pitch along the blade by a similar amount. In some embodiments, in step 915 a different twist may be induced and the blade may be rotated. Various distributions of twist and/or pitch may be used, such as those shown in FIG. 8B.

[0176] The method 900 then moves to step 920 wherein the aircraft 100 takes off vertically. The step 920 may include, for example, rotating the main rotor 210 at a speed and pitch which creates sufficient lift for the aircraft 100 to take off vertically. During vertical flight, the blades may act like an efficient rotor, which may provide cyclic or collective control. Step 920 may include the vertical flight phase, as described herein. The aircraft 100 in step 920 may have a trajectory where the longitudinal axis of the aircraft 100 may or may not be aligned with a geographic vertical or line of action of gravity.

[0177] The method 900 then moves to step 925 wherein the aircraft 100 is rotated for horizontal flight. Step 925 may include rotating the aircraft using cyclic control, falling or diving to rotate to horizontal using control surfaces, etc. Step 925 may include, for example, rotating the aircraft 100 such that the longitudinal axis of the aircraft 100 is generally horizontal. The aircraft 100 may have trajectories where the Z axis of the aircraft 100 may or may not be aligned with a geographic vertical or line of action of gravity.

[0178] The method 900 then moves to step 930 wherein a second electric current is supplied to the piezo element 740 of the blade 720 of the aircraft 100. Step 930 may include, for example, applying the second electric current from an energy store to the piezo element 740. The second electric current may induce a deformation of the piezo element 740 in step 930. In some embodiments, the second electric current is greater than or less than the first electric current.

[0179] The method 900 then moves to step 935 wherein the shape of the blade 720 is changed to a second shape having a second twist. The step 935 may include, for example, the piezo element 740 embedded in the blade 720 inducing a twist in the blade 720, thus altering the blade structure. In some embodiments, the second twist includes a distribution of pitches along the span of the blade that is greater than a first distribution of pitches along the span of the blade due to a first twist, such as the first twist induced in step 915. In some embodiments, the second twist causes the pitch to be about 20 degrees to 30 degrees greater at one or more locations along the span of the blade compared to a distribution of pitches due to a first twist. The second twist may non-uniformly or uniformly change the pitch along the longitudinal or spanwise direction of the blade. In some embodiments, in step 935 the entire blade may be rotated, for example about the blades longitudinal axis, to uniformly change the pitch along the blade by a similar amount. In some embodiments, in step 915 a different twist may be induced and the blade may be rotated. Various distributions of twist and/or pitch may be used, such as those shown in FIG. 8B.

[0180] The method 900 then moves to step 940 wherein the aircraft 100 flies horizontally. The step 940 may include the horizontal flight phase as described herein, for example where lift to the aircraft 100 is provided primarily from the wings and not primarily from the propellers. The speed of rotation of the blade 720 may also be changed in step 935, for example decreased.

[0181] The method 900 then moves to step 945 wherein the aircraft 100 is rotated for vertical landing. The step 945 may include, for example, rotating the aircraft 100 so that the longitudinal axis of the aircraft 100 is in the generally vertical direction with the tails 180, 190 pointed down to the ground. The aircraft 100 may be rotated in step 945 using cyclic controls, control surfaces, by flying upward, etc.

[0182] The method 900 then moves to step 950 wherein a third electric current is supplied to the piezo element 740 of the blade 720 of the aircraft 100. Step 950 may include, for example, applying the third electric current from an energy store to the piezo element 740. In some embodiments, the third electric current may be the same or substantially similar to the first electric current. In some embodiments, the third electric current may be greater than or less than the first electric current. The third electric current may induce a deformation of the piezo element 740 in step 930.

[0183] The method 900 then moves to step 955 wherein the shape of the blade 720 is changed to a third shape having a third twist. Step 955 may include, for example, the piezo element 740 embedded in the blade 720 inducing a twist in the blade 720, thus altering the blade structure. The third twist may be the same or substantially similar to the first twist. The third twist may be greater than or less than the first twist. By“less or greater” it is meant a pitch at a particular spanwise station is less or greater than the pitch at that same spanwise station after inducing a different twist. The difference between the first pitch corresponding to the first twist and the third pitch corresponding to the third twist, at a particular spanwise station, may be about 2 degrees or less. The third twist may include a third pitch distribution that is suitable for landing. The third twist may non-uniformly or uniformly change the pitch along the longitudinal or spanwise direction of the blade. In some embodiments, in step 955 the entire blade may be rotated, for example about the blades longitudinal axis, to uniformly change the pitch along the blade by a similar amount. In some embodiments, in step 955 a different twist may be induced and the blade may be rotated. Various distributions of twist and/or pitch may be used, such as those shown in FIG. 8B.

[0184] The method 900 then moves to step 960 wherein the aircraft 100 lands verticallywith the nose up. Step 960 may include, for example, the longitudinal axis of the aircraft 100 in the generally vertical direction with the tails 180, 190 and/or other structures pointed down and touching down on the ground. The nose may be pointing vertically or off- vertically.

[0185] FIG. 9 is a flowchart showing an embodiment of a method 1000 for operating the hybrid propulsion system 105 and the shape-changing aircraft blade 720 of the tail-sitter aircraft 100 in flight. While described with respect to a particular aircraft, hybrid propulsion system and blades, the method 1000 may be performed with the various aircraft, hybrid propulsion systems and blades described herein, such as the aircraft 100, hybrid propulsion system 105 and blades 212, 214, 216, 720, 720’, the aircraft 100 having wingtip rotor assemblies 230, etc. The method 1000, or portions thereof, may be employed with other methods described herein, for example the methods 1710 and 1730 shown and described with respect to FIGS. 14A-14B.

[0186] The method 1000 begins with step 1010 wherein the shape of the blade 720 is changed to a first shape having a first twist. Step 1010 may be the same or similar to step 915 of the method 900 in FIG. 8.

[0187] The method 1000 then moves to step 1020 wherein a high power is supplied for a short duration from a hybrid propulsion system. Step 1020 may include, for example, providing high or peak power from the hybrid propulsion system 105 for takeoff. The power provided from the hybrid propulsion system 105 for takeoff is expected to be higher than the power provided from the hybrid propulsion system 105 for horizontal flight. The short duration of power provided for takeoff is expected to be shorter than the duration of power provided for horizontal flight. In some embodiments, the hybrid propulsion system 105 may provide the high power using the prime power subsystem 300A, 300B, 300C, 300D or 300E, the electrical power subsystem 200A, 200B, 200C, 200D or 200E, or a combination of the prime and electrical power subsystems. The high power provided is sufficient to lift the aircraft 100 from the ground. The short duration is sufficient to lift the aircraft 100 to a desired height.

[0188] The method 1000 moves to step 1030 wherein the shape of the blade 720 is changed to a second shape having a second twist. Step 1030 may be the same or similar to step 935 of the method 900 in FIG. 8.

[0189] The method 1000 then moves to step 1040 wherein a low power is supplied for a long duration from the hybrid propulsion system 105. The step 1040 may include, for example, providing low power from the hybrid propulsion system 105 for horizontal, sustained flight. The low power provided from the hybrid propulsion system 105 for horizontal flight is expected to be lower than the high power provided from the hybrid propulsion system 105 for takeoff. The long duration of power provided for horizontal flight is expected to be longer than the short duration of power provided for takeoff. In some embodiments, the hybrid propulsion system 105 may provide the low power using the prime power subsystem 300 A, 300B, 300C, 300D or 300E, the electrical power subsystem 200A, 200B, 200C, 200D or 200E, or a combination of the prime and electrical power subsystems. The low power provided is sufficient to keep the aircraft 100 aloft. The long duration is sufficient for the aircraft to go a desired distance or desired amount of time.

[0190] The method 1000 then moves to step 1050 wherein the shape of the blade is changed to a third shape having a third twist. Step 1050 may be the same or similar to step 955 of the method 900 in FIG. 8. The method 1000 then moves to step 1060 wherein a high power is supplied for a short duration from the hybrid propulsion system. Step 1060 may be similar to step 1020. Step 1060 may include, for example, providing high power from the hybrid propulsion system 105 for landing. The high power provided from the hybrid propulsion system 105 for landing is expected to be higher than the low power provided from the hybrid propulsion system 105 for horizontal flight. The short duration of power provided for landing is expected to be shorter than the long duration of power provided for horizontal flight. In some embodiments, the hybrid propulsion system 105 may provide the high power using the prime power subsystem 300 A, 300B, 300C, 300D or 300E, the electrical power subsystem 200A, 200B, 200C, 200D or 200E, or a combination of the prime and electrical power subsystems. The high power provided is sufficient to lower the aircraft 100 to the ground. The short duration is sufficient for the aircraft 100 to land. The high power provided and short duration for landing may be the same or substantially similar to the high power provided and short duration for takeoff. In some embodiments, the high power provided and/or short duration for landing may be less than the high power and/or short duration for takeoff. In some embodiments, the high power provided and/or short duration for landing may be greater than the high power and/or short duration for takeoff.

[0191] FIGS. 10A, 10B and 10C are side, top and front views, respectively, of another embodiment of an aircraft 100’ that may use the wingtip rotor assemblies 230. The aircraft 100’ may also use the shape-changing blades and/or hybrid propulsion systems described herein. The aircraft 100’ may have the same or similar features and/or functionalities as the aircraft 100 described herein, and vice versa.

[0192] The aircraft 100’ includes a hybrid propulsion system 105’. The hybrid propulsion system 105’ may have the same or similar features and/or functionalities as the hybrid propulsion systems 105, 400 or 400A. The hybrid propulsion system 105’ includes an electric power subsystem 200’ and a prime power subsystem 300’. The electric power subsystem 200’ and prime power subsystem 300’ may have the same or similar features and/or functionalities as, respectively, the other electric and prime power subsystems described herein. The electric power subsystem 200’ may be coupled with a main rotor 210’. The bodies of the electric power subsystem 200’, the prime power subsystem 300’ and/or other structural portions of the aircraft 100’ may be part of the fuselage of the aircraft 100’

[0193] The aircraft 100’ may include one main rotor 210’. The main rotor 210’ may have the same or similar features and/or functionalities as the main rotor 210, and vice versa. The main rotor 210’ may be operatively connected to the electric power subsystem 200’. The main rotor 210’ and 210” provide forward propulsive forces to the aircraft 100’, i.e. lift during the vertical takeoff and landing flight phases, and thrust during the horizontal flight phase. The blades of the main rotor 210’ may include piezo elements, such as the piezo elements 740, to change shape, as described herein.

[0194] The aircraft 100’ includes two vertical tails 180’, 190’. The tails 180’, 190’ are angled, vertical segments extending from a rear portion of a center fuselage portion of the aircraft 100’. The tails 180’ provide a support and landing structure for the aircraft 100’. The tails 180’ may extend rearward and outward in the X-Z plane or a plane parallel thereto and at an angle with respect to the X axis and Z axis. The tails 180’ may the same or similar features and/or functionalities as the other tails described herein, for example the tails 180, 190, etc. The aircraft 100’ may include supports 191 located outboard of the longitudinal axis. The supports 191 may be inboard from the wingtip regions 122, 132. The supports 191 may provide support and balance to the aircraft 100’ while on the ground, during takeoff, during landing, etc. The vertical tails 180’ and the supports 190’ provide four points of contact with the ground and a larger footprint for the aircraft 100’ for greater stability. [0195] The aircraft 100’ may be operated using the various methods described herein, for example the methods 500, 1710, 1730 and others. The aircraft 100’ may include the various blade assemblies described herein, for example the blade assembly 700 and/or blade 720 or 720’.

[0196] FIGS. 11A-11C are various views of another embodiment of an aircraft 101 having rotatable wingtip rotor assemblies 230A-D. The aircraft 101 includes wings 119A, 119B, 119C, and 119D. The wings may be in the X-wing configuration shown, where each wing is angled with respect to the X-Y plane. The wings 119A and 119B are angled upward in the +Z direction relative to the X-Y plane, and the wings 119C and 119D are angled downward in the -Z direction relative to the X-Y plane. The angle between upper and lower wings may be various amounts, for example from about 3 degrees to about 45 degrees, from about 15 degrees to about 45 degrees, from about 15 degrees to about 30 degrees, or other ranges or angular amounts.

[0197] The aircraft 101 may include wingtip rotor assemblies 230A-D at the wingtip regions of respective wings 119A-D. FIGS. 11 A-l 1B show the assemblies 230A and 230B oriented to provide thrust vectors 1 (TV 1) and 2 (TV2) in the +Z direction and the assemblies 230C and 230D oriented to provide thrust vectors 3 (TV 3) and 4 (TV 4) in the -Z direction. The assemblies 230A-D may be pushers or pullers. FIG. 11C shows the assemblies 230A-D oriented to provide thrust vectors 1-4 in the +X direction.

[0198] The aircraft 101 may include one or more support structures 160A-B and 170A-B. As shown there are four support structures, 160A-B, 170A-B each attached to a respective wing 119A-D. The support structures 160A-B, 170A-B may support the aircraft for takeoff and after landing with the X-axis oriented vertically or approximately vertically.

[0199] FIGS. 12A-12C are various views of another embodiment of an aircraft 102 having rotatable wingtip rotor assemblies 230A-D. The aircraft 102 may have the same or similar features and/or functionalities as the aircraft 101 of FIGS. 11A-11C, except as otherwise noted. The aircraft 102 may include the rotatable wingtip rotor assemblies 230A-D.

[0200] The aircraft 102 may additionally include one or more fixed inboard rotor assemblies 210A-H. As shown, there may be eight inboard assemblies 210A-H. In some embodiments, there may be one, two, three, four, five, six, seven, nine or more inboard assemblies 210A-210H. There may be one or more assemblies 210A-210D on the wing 119C, and one or more of the assemblies 210E-210H on the wing 119D. As shown, there may be four assemblies 210A-210D on the wing 119C, and four assemblies 210E-210H on the wing 119D. The assemblies 210A-210H may be statically mounted or fixed, e.g. not be rotatable about a lateral axis. The assemblies 210A-210H may be located on the lower wings 119C and 119D as shown. The assemblies 210A-210H may be located, in addition or alternatively, on the upper wings 119A, 119B. The assemblies 210A-210H may be pusher or puller props. The assemblies 210A-210H may rotate in opposite directions to provide counter torque. For example, the assemblies 210A-D may rotate a first direction and assemblies 210E-H may rotate a second direction that is opposite to the first direction. The assemblies 210A-210H may be incorporated on the other aircraft described herein, for example the aircraft 100 of FIGS. 1A- 1C.

[0201] The aircraft 102 is shown in FIGS. 12A-12C with the wingtip assemblies 230 A and 230B oriented to provide thrust in a +Z direction, and the wingtip assemblies 230C and 230D oriented to provide thrust in a -Z direction. The assemblies 230A-D may be pushers or pullers. The assemblies 230A-D may rotate about respective lateral axes (FA) 1-4 as indicated. The assemblies 230A and 230B may rotate the entire body or nacelle as shown. The assemblies 230C and 230D may rotate part of the body or nacelle, such as a nose cap, as shown.

[0202] The aircraft 102 may include one or more inboard rotor assemblies 215 A- D. The assemblies 215A, 2l5Bmay be located within the wing 119B and the assemblies 215C, 215D may be located within the wing 119A. The assemblies 215A-215D may be located between leading and trailing edges of the wings. There may be openings in the wings where the assemblies 215A-215D are located, for example circular openings, vented openings, etc. The assemblies 215A-215D may be embedded with rotors that spin to provide thrust and counter torque depending on the orientation. The assemblies 215 A-215D may also rotate about an axis that is at an angle to the respective rotor’ s spin axis. For example, the assemblies 215C, 215D may rotate about the lateral axis (FA) 5 as shown in FIG. 12B. The assemblies 215A, 215B may rotate about the lateral axis (FA) 6 as shown in FIG. 12B. The assemblies 215A-D may rotate about FA 1 or 2 or axes parallel thereto. The assemblies 215A-D may rotate about a chord of the aircraft for example to angle the thrust vector in the ZY plane. In some embodiments, the assemblies 215A-D may be fixed and not rotatable. The assemblies 215 A- D may provide control of the aircraft, such as thrust, counter torque, other attitude control, etc. The assemblies 215A-D may additionally or alternatively be located on the lower wings 119C and 119D. There may be two assemblies 215A-D on each wing as shown, or one, three or more assemblies 215A-215D on one or more of the wings.

[0203] FIG. 13 is a perspective view of another embodiment of an aircraft 103 having rotatable wingtip rotor assemblies 230A, 23 OB and fixed inboard rotor assemblies 210A, 21 OB. The aircraft 103 may have the same or similar features and/or functionalities as the other aircraft described herein, such as the aircraft 100, etc. The rotor assemblies 210A may be located under the right wing 119 A, and the rotor assemblies 210B may be located under the left wing 119B. The assemblies 210A and 210B may be located below, above, or within their respective wings 119A, 119B, or combinations thereof. There may be twelve assemblies 210A, 210B as shown, or fewer or greater than twelve assemblies 230A.

[0204] In some embodiments, counter torque from the wingtip rotor assemblies 230 A, 230B may not be needed, for example where there is an even number of similar main rotors 210. In such cases, the wingtip rotor assemblies 230 may still be employed for reducing drag, enhancing thrust, etc., and the assemblies 230 may still be rotatable, for example to provide attitude control, etc.

[0205] Other types of aircraft besides those shown and described herein may utilize the various features described herein. For example, the wingtip rotor assemblies 230 may be used with various other aircraft types and with modifications to the aircraft described herein. For instance, the rotor assemblies 230 may be used with aircraft having two wings that are swept forward or backward, or angled upward or downward, with various shapes or planforms, etc. For instance, the rotor assemblies 230 may be used with X-wing configurations of various shapes and sizes, or with more than four wings, etc. Further, there may be more than one rotor assembly 230 for each wing. For instance, there may be two or more rotor assemblies 230 for each wing of an aircraft. The rotor assemblies 230 may be at the wingtip regions or inboard. The rotor assemblies 230 may be both at the wingtip regions and located inboard on a single wing.

[0206] As further example, the rotor assemblies 230 may be used with aircraft having mismatched propeller and rotor torques on the centerline, such as a“push pull” aircraft configuration. There could be a rotor on the front of this type of aircraft if it was a tail-sitter. A large top rotor and small prop on the centerline may cause a torque imbalance that would benefit from wing-tip mounted, counter-torque rotor assemblies 230.

[0207] In some embodiments, the rotor assemblies 230 may not be located at the wingtip regions. For example, the rotor assemblies 230 may be mounted such that they are internal to the wing and can pivot, such as the rotor assemblies 215A-D shown in FIGS. 12A- 12C. The assemblies 230 may be mounted within an opening of a wing and pivot within the opening. This embodiment may be used with tiltrotor aircraft. In some embodiments, the rotor assemblies 230 may be located within the wing as well as at the wingtip regions. Thus, various different types of aircraft and aircraft configurations besides those explicitly shown or described herein may employ the rotatable rotor assemblies 230 and other features described herein.

[0208] FIG. 14A is a flowchart showing an embodiment of a method 1710 for a VTOL aircraft to takeoff using rotatable wingtip rotor assemblies. The method 1710 may be employed by the various aircraft and systems described herein, such as those described with respect to FIGS. 1A-13. For example, the aircraft 100 and wingtip rotor assemblies 230 may be used.

[0209] The method 1710 may begin with step 1712 wherein the wingtip rotor assemblies are oriented for counter-torque. The assemblies 230 may be oriented to provide thrust vectors in opposite Z-directions, for example the orientation shown in FIG. 1B. The assemblies 230 may be rotated using the systems 401 A or 401 B and/or the mechanisms and features shown in FIGS. 1H-1L. The aircraft in step 1712 may be oriented for vertical takeoff. The aircraft 100 may be oriented vertically on its tail for takeoff. Aircraft besides tail-sitters may be used as well. The method 1730 next moves to step 1714 wherein the main rotor(s) and wingtip rotor assemblies are powered. The rotors may be powered using the various systems described herein, such as the systems shown in FIGS. 1D-1E or 4. The method 1710 next moves to step 1716 wherein vertical thrust is provided by the main rotor(s) and counter torque is provided by the wingtip rotor assemblies. The method 1710 then moves to step 1718 wherein the aircraft takes off vertically. The main rotors may provide the upward thrust. The wingtip rotors may provide counter torque to maintain or control rotation of the aircraft about a vertical axis (e.g. , the longitudinal X axis). The wingtip rotors may also provide some upward thrust in this step. [0210] The method 1710 next moves to step 1720 where the aircraft is rotated for forward flight. The wingtip rotors may rotate about lateral axes to rotate the aircraft in step 1720. In addition or alternatively, the main rotor 210 may pivot and/or have shape-changing blades that alter the blade’s shape or configuration to rotate the aircraft in step 1720. The aircraft may actuate wing control surfaces in step 1720. The method 1710 next moves to step 1722 wherein the wingtip rotor assemblies are rotated to provide forward thrust. The assemblies may be rotated in step 1722 using the systems 401 A or 401 B and/or the mechanisms and features shown in FIGS. 1H-1L.

[0211] FIG. 14B is a flowchart showing an embodiment of a method 1730 for a VTOL aircraft to land using rotatable wingtip rotor assemblies. The method 1730 may be employed by the various aircraft and systems described herein, such as those described with respect to FIGS. 1A-13. The method 1730 may begin with step 1732 wherein an aircraft is oriented, for example rotated, for vertical flight. The aircraft may be in forward flight and then oriented for vertical flight in step 1732. The aircraft 100 may be oriented vertically for landing on its tail. Aircraft besides tail-sitters may be used as well. The wingtip rotors may rotate about lateral axes to rotate the aircraft in step 1732. In addition or alternatively, the main rotor 210 may pivot and/or have shape-changing blades that alter the blade’s shape or configuration to rotate the aircraft in step 1732. The aircraft may actuate wing control surfaces in step 1732.

[0212] The method 1730 next moves to step 1734 wherein the wingtip rotor assemblies are oriented for counter-torque. The assemblies 230 may be oriented to provide thrust vectors in opposite Z-directions, for example the orientation shown in FIG. 1B. The assemblies 230 may be rotated using the systems 401 A or 401 B and/or the mechanisms and features shown in FIGS. 1H-1L. The method 1730 then moves to step 1736, which may be the same or similar as step 1716 of the method 1710. The method 1730 then moves to step 1738 wherein the aircraft lands, for example on its tail or otherwise lands vertically. The wingtip assemblies 230 in steps 1736 and/or 1738 may be oriented to provide purely or mostly counter torque or the assemblies 230 may be oriented to have a thrust component in the upward direction (e.g., the +X direction for a tail sitter).

[0213] While the above detailed description has shown, described, and pointed out novel features of the invention as applied to various embodiments, it will be understood that various omissions, substitutions, and changes in the form and details of the device or process illustrated may be made by those skilled in the art without departing from the spirit of the invention. As will be recognized, the present invention may be embodied within a form that does not provide all of the features and benefits set forth herein, as some features may be used or practiced separately from others. The scope of the invention is indicated by the appended claims rather than by the foregoing description. All changes which come within the meaning and range of equivalency of the claims are to be embraced within their scope.

[0214] The foregoing description details certain embodiments of the systems, devices, and methods disclosed herein. It will be appreciated, however, that no matter how detailed the foregoing appears in text, the systems, devices, and methods may be practiced in many ways. As is also stated above, it should be noted that the use of particular terminology when describing certain features or aspects of the invention should not be taken to imply that the terminology is being re-defined herein to be restricted to including any specific characteristics of the features or aspects of the technology with which that terminology is associated.

[0215] It will be appreciated by those skilled in the art that various modifications and changes may be made without departing from the scope of the described technology. Such modifications and changes are intended to fall within the scope of the embodiments. It will also be appreciated by those of skill in the art that parts included in one embodiment are interchangeable with other embodiments; one or more parts from a depicted embodiment may be included with other depicted embodiments in any combination. For example, any of the various components described herein and/or depicted in the Figures may be combined, interchanged or excluded from other embodiments.

[0216] The flow chart sequences are illustrative only. A person of skill in the art will understand that the steps, decisions, and processes embodied in the flowcharts described herein may be performed in an order other than that described herein. Thus, the particular flowcharts and descriptions are not intended to limit the associated processes to being performed in the specific order described.

[0217] With respect to the use of substantially any plural and/or singular terms herein, those having skill in the art may translate from the plural to the singular and/or from the singular to the plural as is appropriate to the context and/or application. The various singular/plural permutations may be expressly set forth herein for sake of clarity. [0218] It will be understood by those within the art that, in general, terms used herein are generally intended as“open” terms (e.g., the term“including” should be interpreted as“including but not limited to,” the term“having” should be interpreted as“having at least,” the term“includes” should be interpreted as“includes but is not limited to,” etc.). It will be further understood by those within the art that if a specific number of an introduced claim recitation is intended, such an intent will be explicitly recited in the claim, and in the absence of such recitation no such intent is present. For example, as an aid to understanding, the following appended claims may contain usage of the introductory phrases“at least one” and “one or more” to introduce claim recitations. However, the use of such phrases should not be construed to imply that the introduction of a claim recitation by the indefinite articles“a” or “an” limits any particular claim containing such introduced claim recitation to embodiments containing only one such recitation, even when the same claim includes the introductory phrases“one or more” or“at least one” and indefinite articles such as“a” or“an” (e.g.,“a” and/or“an” should typically be interpreted to mean“at least one” or“one or more”); the same holds true for the use of definite articles used to introduce claim recitations. In addition, even if a specific number of an introduced claim recitation is explicitly recited, those skilled in the art will recognize that such recitation should typically be interpreted to mean at least the recited number (e.g., the bare recitation of“two recitations,” without other modifiers, typically means at least two recitations, or two or more recitations). Furthermore, in those instances where a convention analogous to“at least one of A, B, and C, etc.” is used, in general such a construction is intended in the sense one having skill in the art would understand the convention (e.g.,“a system having at least one of A, B, and C” would include but not be limited to systems that have A alone, B alone, C alone, A and B together, A and C together, B and C together, and/or A, B, and C together, etc.). In those instances where a convention analogous to“at least one of A, B, or C, etc.” is used, in general such a construction is intended in the sense one having skill in the art would understand the convention (e.g.,“a system having at least one of A, B, or C” would include but not be limited to systems that have A alone, B alone, C alone, A and B together, A and C together, B and C together, and/or A, B, and C together, etc.). It will be further understood by those within the art that virtually any disjunctive word and/or phrase presenting two or more alternative terms, whether in the description, claims, or drawings, should be understood to contemplate the possibilities of including one of the terms, either of the terms, or both terms. For example, the phrase“A or B” will be understood to include the possibilities of“A” or“B” or“A and B.”

[0219] All references cited herein are incorporated herein by reference in their entirety. To the extent publications and patents or patent applications incorporated by reference contradict the disclosure contained in the specification, the specification is intended to supersede and/or take precedence over any such contradictory material.

[0220] The term“comprising” as used herein is synonymous with“including,” “containing,” or“characterized by,” and is inclusive or open-ended and does not exclude additional, unrecited elements or method steps.

[0221] All numbers expressing quantities of ingredients, reaction conditions, and so forth used in the specification and claims are to be understood as being modified in all instances by the term“about.” Accordingly, unless indicated to the contrary, the numerical parameters set forth in the specification and attached claims are approximations that may vary depending upon the desired properties sought to be obtained by the present invention. At the very least, and not as an attempt to limit the application of the doctrine of equivalents to the scope of the claims, each numerical parameter should be construed in light of the number of significant digits and ordinary rounding approaches.

[0222] The above description discloses several methods and materials of the present invention. This invention is susceptible to modifications in the methods and materials, as well as alterations in the fabrication methods and equipment. Such modifications will become apparent to those skilled in the art from a consideration of this disclosure or practice of the invention disclosed herein. Consequently, it is not intended that this invention be limited to the specific embodiments disclosed herein, but that it cover all modifications and alternatives coming within the true scope and spirit of the invention as embodied in the attached claims.