Stop-rotor aircraft

Abstract

An aircraft includes a base frame, a main rotor assembly including at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, the at least one wing defining a center of pressure, and a component adjustment assembly arranged on the base frame assembly and configured to move at least one component of the aircraft relative to the base frame assembly to one or more of move the center of pressure of the at least one wing relative to a center of gravity of the aircraft or move the center of gravity relative to the center of pressure.

Description

Attorney Docket No.: MIT 25110 PCT | 88212-425635 STOP-ROTOR AIRCRAFT CROSS REFERENCE TO RELATED APPLICATION

[0001] The present disclosure claims priority to and the benefit of U.S. Provisional Patent Application No.63 / 687,972, entitled “STOP ROTOR AIRCRAFT,” filed August 28, 2024, the disclosure of which is incorporated by reference herein in its entirety. FIELD

[0002] The present disclosure relates to aircraft, and in particular, aircraft capable of vertical takeoff and landing. BACKGROUND

[0003] Vertical takeoff and landing (VTOL) aircraft have long been explored, driven by desires to operate such aircraft in confined spaces, remote locations, and challenging environments where traditional runways are impractical or unavailable. Since their development, VTOL aircraft have emerged in the military, civil, and commercial sectors. In military applications, VTOL platforms provide rapid deployment, close air support, and reconnaissance due to the increase agility and responsiveness offered by the transitionary system. In the civil and commercial sectors, VTOL aircraft offer promising applications for environmental monitoring and tracking, urban air mobility, and search and rescue. In essence, VTOL aircraft offer a path for uncrewed aerial vehicles (UAV) to become omnipresent across military, commercial, and civil sectors.

[0004] Helicopters emerged as a suitable candidate as a VTOL aircraft due to their highly maneuverable flight and hover efficiency, offering a versatile alternative to traditional fixed- wing aircraft and access to remote areas. However, the relatively low speeds, limited range, and poor cruise performance of helicopters can limit missions to hover-based tasks or short point-to-point missions. As such, researchers have continued to search for an alternative VTOL aircraft capable of bridging the forward flight performance of traditional airplanes with the hover and maneuverability performance of helicopters.

[0005] Numerous VTOL platforms have been developed that aim to address efficiency limitations across flight modes. Some include tiltrotor and thrust-vectored formulations. Despite advancements in VTOL technologies, many developed platforms still favor hover orAttorney Docket No.: MIT 25110 PCT | 88212-425635 forward flight efficiency, rather than optimizing across both flight modes. This limitation commonly stems from the trade-offs inherent to aerodynamic design, propulsion systems, and control mechanisms. For example, optimizing a VTOL aircraft for efficient hover often involves the use of large rotor systems or ducted fans, which may result in increased weight and drag during forward flight, consequently compromising forward flight efficiency. Conversely, aircraft optimized for forward flight utilize large passive wing areas and small thrust propellers, inhibiting the efficiency in VTOL.

[0006] Numerous VTOL platforms have been developed that aim to address efficiency limitations across flight modes. Some include tiltrotor and thrust-vectored formulations. Despite advancements in VTOL technologies, many developed platforms still favor hover or forward flight efficiency, rather than optimizing across both flight modes. This limitation commonly stems from the trade-offs inherent to aerodynamic design, propulsion systems, and control mechanisms. For example, optimizing a VTOL aircraft for efficient hover often involves the use of large rotor systems or ducted fans, which may result in increased weight and drag during forward flight, consequently compromising forward flight efficiency. Conversely, aircraft optimized for forward flight utilize large passive wing areas and small thrust propellers, inhibiting the efficiency in VTOL.

[0007] Certain point-to-point missions, such as urban air mobility taxi services or shuttle flights between offshore platforms and onshore facilities, may not necessitate prolonged hover times. These point-to-point missions can take on a mission profile including a short period of time is spent in VTOL / hover mode, and a majority of flight time is spent in forward flight. As such, the efficiency of VTOL / hover does not influence the overall mission profile. However, more VTOL missions are emerging that require the aircraft to operate equally, or near equally, in both hover and forward flight, such as search and rescue missions, monitoring, and mapping. Search and rescue missions have several periods of switching back and forth between VTOL / hover and forward flight as the aircraft traverses and then stops to observe. As such, developing an aircraft that can operate efficiently across both hover and forward flight would be advantageous.

[0008] The stop-rotor aircraft has emerged as a promising solution for achieving improved efficiency in both VTOL and forward flight operation. During hover, takeoff, and landing, the central lifting surface functions as a rotor, generating vertical thrust via the rotation of the wing. Once airborne, the central lifting surface is stopped and fixed in place, where itAttorney Docket No.: MIT 25110 PCT | 88212-425635 generates passive lift in forward flight similar to a conventional fixed-wing aircraft. The dual usage of the blade aims to optimize the aerodynamic properties across both flight modes. As such, stop-rotor aircraft are typically characterized by certain advantages which can include, for example, enhanced efficiency, enhanced speed and range, enhanced payload capacity, and enhanced versatility.

[0009] Specifically, regarding efficiency, the dual usage of the blade optimizes aerodynamic performance in each flight mode: maximal VTOL capability is achieved with a large rotor area, similar to that of a helicopter, while optimal cruise efficiency is achieved with a large lifting surface, resembling that of a conventional aircraft. This dual usage of the central lifting surface ensures optimized performance in both flight modes without the need for additional hardware such as additional or oversized power plants. Regarding speed and range, compared to a helicopter blade that is subject to advancing blade compressibility losses and retreating blade stall in forward flight, the passive lifting surface removes blade stall and energy usage on the rotor during forward flight. Furthermore, the minimal design of the stop-rotor aircraft will reduce drag in forward flight (i.e., due to the lack of additional or oversized turbofan engines), thereby reducing energy usage for a given speed and in turn improving the range and maximum speed.

[0010] Regarding payload capacity, in both flight modes, the stop-rotor aircraft will have similar payload capacity to that of the analog system (i.e., helicopter in hover or conventional aircraft in forward flight). In hover, the large rotor area enhances the lift capacity. In forward flight, the lack of oversized turbofans and propulsion systems translates to increased payload capacity. Regarding versatility, stop-rotor aircraft combine the VTOL capabilities of helicopters with the speed and efficiency of fixed-wing aircraft. This versatility allows them to operate in diverse environments, including urban areas, remote locations, and confined spaces where traditional aircraft and cruise preferred aircraft may have limited access.

[0011] Stop-rotor aircraft have typically been limited by three main challenges. First, the airfoil directionality between VTOL and forward flight modes differ. In VTOL, as the rotor spins, the airfoils should be on opposite sides of the wing, as shown in “VTOL” in FIG 2. In forward flight, the airfoils should be on the same side of the wing to generate lift, as shown in “Forward Flight” in FIG.2. The conflicting airfoil directionality can add either mechanical or controller complexity, and / or can hinder the overall aerodynamic performance.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0012] Second, an optimal location of the center of pressure across VTOL and forward flight can differ: in VTOL, the center of pressure, also referred to as the axis of rotation, ideally passes through the center of gravity and the center of counter torque, as shown in “VTOL” in FIG.3. So long as the center of rotation is aligned with the center of gravity and the center of counter torque, the system is passively stable in yaw and a controller associated with counter torque can be used to stabilize any disturbances. In forward flight, the center of pressure should be aft the center of gravity to ensure passive pitch stability, as shown in “Forward Flight” in FIG.3. A disturbance that pitches the aircraft nose up is corrected by a subsequent increase in the lift force and decrease in down force, while a pitch down disturbance is corrected by a passive decrease in lift force and increase in down force.

[0013] Finally, the transition between rotational and translational lift generation can lead to a temporary loss of lift which makes altitude control challenging. As a result, system complexity needs to increase, either mechanically or in the controller, to ensure the aircraft can maintain stable altitude across the transition.

[0014] Accordingly, there is a need for an optimized stop-rotor aircraft that addresses the above-described challenges and shortcomings. SUMMARY

[0015] This Summary introduces a selection of concepts in simplified form that are described further below in the Detailed Description. This Summary does not necessarily identify key or essential features, nor does it limit the scope of the claimed subject matter.

[0016] The stop-rotor aircrafts described herein provide a potential solution to the above- described shortcomings of stop-rotor aircraft technology for VTOL operations. Specifically, the disclosed stop-rotor aircrafts provide a novel stop-rotor platform capable of bi-directional VTOL to forward flight transition. The stop-rotor aircrafts address the above-described challenges, in particular the difference in airfoil directionality between VTOL and forward flight modes, the difference in center of pressure across VTOL and forward flight modes, and the transition between rotational and translational lift generation leading to a temporary loss of lift, and provide improved aircrafts and associated control schema that mitigate and potentially eliminate the effects of these challenges.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0017] A stop-rotor aircraft according to a first aspect of the present disclosure includes a base frame assembly, a main rotor assembly including at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, the at least one wing defining a center of pressure, and a component adjustment assembly arranged on the base frame assembly and configured to move at least one component of the aircraft relative to the base frame assembly to one or more of move the center of pressure of the at least one wing relative to a center of gravity of the aircraft or move the center of gravity relative to the center of pressure.

[0018] In some embodiments, the at least one component of the aircraft that can be movable by the component adjustment assembly includes the main rotor assembly. The component adjustment assembly can be configured to move the main rotor assembly in a forward-aft direction to move the center of pressure of the at least one wing in the forward- aft direction. The component adjustment assembly can include a carriage configured to move in the forward-aft direction relative to the base frame assembly, and the main rotor assembly can be coupled to the carriage of the component adjustment assembly such that movement of the carriage relative to the base frame assembly moves the center of pressure of the at least one wing forward and aft relative to the center of gravity of the aircraft.

[0019] In at least some embodiments, the carriage can be movable between a forward position for VTOL flight of the aircraft in which the center of pressure of the at least one wing is approximately aligned with the center of gravity of the aircraft in the forward-aft direction and an aft position for forward flight of the aircraft in which the center of pressure of the at least one wing is aft of the center of gravity of the aircraft in the forward-aft direction.

[0020] In some such embodiments, the component adjustment assembly can further include a linear rail arranged on the base frame assembly that extends in the forward-aft direction, and the carriage is slidably movable on the linear rail. The component adjustment assembly can further include a linear actuator and an actuator rod coupled to the carriage, and the linear actuator can be configured to move the actuator rod toward and away from the linear actuator to move the carriage toward and away from the linear actuator along the linear rail.

[0021] In some embodiments, the aircraft can further include a motor assembly including a motor coupled to a central housing of the main rotor assembly from which the at least oneAttorney Docket No.: MIT 25110 PCT | 88212-425635 wing extends, the motor configured to rotate the main rotor assembly, the motor assembly being arranged on the carriage of the component adjustment assembly such that the motor assembly and the main rotor assembly are configured to be moved relative to the base frame assembly via movement of the carriage relative to the base frame assembly. The at least one wing can be rotatable relative to the central housing between a VTOL position in which a leading edge of the at least one wing faces a direction of rotation of the main rotor assembly about the axis and a forward flight position in which the leading edge faces a forward direction of the aircraft. The at least one wing can include a first wing that is fixed relative to the central housing and a second wing that is rotatable relative to the central housing, and the motor and first and second wings can be configured to be rotationally locked in forward flight. The leading edges of the first and second wings can face opposite directions in the VTOL position of the second wing, and the leading edges of the first and second wings can face the forward direction in the forward flight position of the second wing.

[0022] In at least some embodiments, the main rotor assembly can further include a transition motor assembly configured to rotate the at least one wing between and including the VTOL and forward flight positions, and the transition motor assembly can include a locking bumper arranged in a fixed position relative to the at least one wing. The transition motor assembly can further include a locking pin configured to be engaged by the locking bumper during rotation of the at least one wing from the VTOL position to the forward flight position, and be configured to be disengaged from the locking bumper during rotation of the at least one wing from the forward flight position to the VTOL position. In certain embodiments, engagement of the locking bumper with the locking pin during rotation of the at least one wing from the VTOL position to the forward flight position can cause the locking pin to enter a hole formed in a rotor plate fixedly disposed on the base frame assembly to rotationally lock the main rotor assembly relative to the base frame assembly. Disengagement of the locking bumper from the locking pin during rotation of the at least one wing from the forward flight position to the VTOL position can cause the locking pin to be removed from a hole formed in a rotor plate fixedly disposed on the base frame assembly to rotationally unlock the main rotor assembly relative to the base frame assembly.

[0023] In some such embodiments, the aircraft may further include a counterbalance assembly including a first counterbalance subassembly having a first counterbalance rotor and extending away from a first side of the base frame assembly and a second counterbalanceAttorney Docket No.: MIT 25110 PCT | 88212-425635 subassembly having a second counterbalance rotor and extending away from a second side of the base frame assembly opposite the first side. The first and second counterbalance rotors can be rotatable between a forward-facing position and an aft-facing position. In some embodiments, in a VTOL configuration, the first counterbalance rotor is in the forward- facing position and the second counterbalance rotor can be in the aft-facing position such that first and second counterbalance rotors face a direction of rotation of the main rotor assembly about the axis such that the first and second counterbalance rotors stabilize a yaw of the aircraft. In some embodiments, in a forward flight configuration, the first counterbalance rotor is in the forward-facing position and the second counterbalance rotor can be in the forward-facing position such that first and second counterbalance rotors face a forward direction of the aircraft to generate forward thrust to the aircraft.

[0024] In certain embodiments, the aircraft may further include a multicopter assembly arranged on the base frame assembly and including a multicopter plate and a plurality of multicopter rotors arranged on the multicopter plate, the plurality of multicopter rotors configured to generate thrust in a downward direction so as to provide lift to the aircraft. The plurality of multicopter rotors can be configured to be rotating in a VTOL mode of the aircraft and in a transition mode of the aircraft between VTOL and forward flight to provide lift to the aircraft, and can be configured to be shut off in a forward flight mode of the aircraft. The plurality of multicopter rotors can include four multicopter rotors. The multicopter plate can be arranged on the base frame assembly above the center of gravity of the aircraft. In certain examples, the multicopter plate can be arranged on the base frame assembly above the center of gravity of the aircraft such that a center of thrust of the plurality of multicopter rotors is arranged above the center of gravity.

[0025] In at least some embodiments, the at least one component of the aircraft that is movable by the component adjustment assembly can include a first component of the aircraft, and the component adjustment assembly is configured to move the first component in a forward-aft direction to move the center of gravity of the aircraft in the forward-aft direction. In some embodiments, the first component can be an electronics component of the aircraft. The component adjustment assembly can include a carriage configured to move in the forward-aft direction relative to the base frame assembly, and the first component can be coupled to the carriage of the component adjustment assembly such that movement of the carriage relative to the base frame assembly moves the center of gravity forward and aftAttorney Docket No.: MIT 25110 PCT | 88212-425635 relative to the center of pressure of the at least one wing. The carriage can be movable between an aft position for VTOL flight of the aircraft in which the center of pressure of the at least one wing is approximately aligned with the center of gravity of the aircraft in the forward-aft direction and a forward position for forward flight of the aircraft in which the center of pressure of the at least one wing is aft of the center of gravity of the aircraft in the forward-aft direction. The component adjustment assembly can further include a linear rail arranged on an underside of the base frame assembly that extends in the forward-aft direction, and the carriage can be slidably movable on the linear rail. The component adjustment assembly can further include a linear actuator and an actuator rod coupled to the carriage, and the linear actuator can be configured to move the actuator rod toward and away from the linear actuator to move the carriage toward and away from the linear actuator along the linear rail.

[0026] A method according to a further aspect of the present disclosure includes providing a base frame assembly of an aircraft, coupling a main rotor assembly of the aircraft to a component adjustment assembly of the aircraft, the main rotor assembly including at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, the at least one wing defining a center of pressure, and arranging the component adjustment assembly on the base frame assembly, the component adjustment assembly being configured to move at least one component of the aircraft relative to the base frame assembly to one or more of move the center of pressure of the at least one wing relative to a center of gravity of the aircraft or move the center of gravity relative to the center of pressure.

[0027] In some embodiments, the at least one component of the aircraft that is movable by the component adjustment assembly can include the main rotor assembly. The component adjustment assembly can be configured to move the main rotor assembly in a forward-aft direction to move the center of pressure of the at least one wing in the forward-aft direction. The component adjustment assembly can include a carriage configured to move in the forward-aft direction relative to the base frame assembly, and the method can further include coupling the main rotor assembly to the carriage of the component adjustment assembly such that movement of the carriage relative to the base frame assembly moves the center of pressure of the at least one wing forward and aft relative to the center of gravity of the aircraft. The carriage can be movable between a forward position for VTOL flight of theAttorney Docket No.: MIT 25110 PCT | 88212-425635 aircraft in which the center of pressure of the at least one wing is approximately aligned with the center of gravity of the aircraft in the forward-aft direction and an aft position for forward flight of the aircraft in which the center of pressure of the at least one wing is aft of the center of gravity of the aircraft in the forward-aft direction.

[0028] In some such embodiments, the method can further include arranging a linear rail of the component adjustment assembly on the base frame assembly, the linear rail extending in the forward-aft direction, and the carriage being slidably movable on the linear rail. The method can further include arranging a linear actuator of the component adjustment assembly on the base frame, the component adjustment assembly further including an actuator rod that is actuatable by the linear actuator, and coupling the actuator rod to the carriage, the linear actuator being configured to move the actuator rod toward and away from the linear actuator to move the carriage toward and away from the linear actuator along the linear rail.

[0029] In at least some embodiments, the method can further include coupling a motor of a motor assembly of the aircraft to a central housing of the main rotor assembly from which the at least one wing extends, the motor being configured to rotate the main rotor assembly, and arranging the motor assembly on the carriage of the component adjustment assembly such that the motor assembly and the main rotor assembly are configured to be moved relative to the base frame assembly via movement of the carriage relative to the base frame assembly. In some embodiments, the at least one wing can be rotatable relative to the central housing between a VTOL position in which a leading edge of the at least one wing faces a direction of rotation of the main rotor assembly about the axis and a forward flight position in which the leading edge faces a forward direction of the aircraft. The at least one wing can include a first wing that is fixed relative to the central housing and a second wing that is rotatable relative to the central housing, and the motor and first and second wings can be configured to be rotationally locked in forward flight. The leading edges of the first and second wings can face opposite directions in the VTOL position of the second wing, and the leading edges of the first and second wings can face the forward direction in the forward flight position of the second wing. The main rotor assembly can further include a transition motor assembly configured to rotate the at least one wing between and including the VTOL and forward flight positions, and the transition motor assembly can include a locking bumper arranged in a fixed position relative to the at least one wing.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0030] In some embodiments, the transition motor assembly can further include a locking pin configured to be engaged by the locking bumper during rotation of the at least one wing from the VTOL position to the forward flight position, and configured to be disengaged from the locking bumper during rotation of the at least one wing from the forward flight position to the VTOL position. Engagement of the locking bumper with the locking pin during rotation of the at least one wing from the VTOL position to the forward flight position can cause the locking pin to enter a hole formed in a rotor plate fixedly disposed on the base frame assembly to rotationally lock the main rotor assembly relative to the base frame assembly. Disengagement of the locking bumper from the locking pin during rotation of the at least one wing from the forward flight position to the VTOL position can cause the locking pin to be removed from a hole formed in a rotor plate fixedly disposed on the base frame assembly to rotationally unlock the main rotor assembly relative to the base frame assembly.

[0031] In some such embodiments, the method can further include arranging a first counterbalance subassembly of a counterbalance assembly of the aircraft on a first side of the base frame assembly such that the first counterbalance subassembly extends away from the first side of the base frame assembly, the first counterbalance subassembly including a first counterbalance rotor, and arranging a second counterbalance subassembly of the counterbalance assembly on a second side of the base frame assembly opposite the first side such that the second counterbalance subassembly extends away from the second side of the base frame assembly, the second counterbalance subassembly including a second counterbalance rotor. The first and second counterbalance rotors can be rotatable between a forward-facing position and an aft-facing position. In some embodiments, in a VTOL configuration, the first counterbalance rotor can be in the forward-facing position and the second counterbalance rotor is in the aft-facing position such that first and second counterbalance rotors face a direction of rotation of the main rotor assembly about the axis such that the first and second counterbalance rotors stabilize a yaw of the aircraft. In some embodiments, in a forward flight configuration, the first counterbalance rotor can be in the forward-facing position and the second counterbalance rotor is in the forward-facing position such that first and second counterbalance rotors face a forward direction of the aircraft to generate forward thrust to the aircraft.

[0032] In at least some embodiments, the method can further include arranging a multicopter assembly of the aircraft on the base frame assembly, the multicopter assemblyAttorney Docket No.: MIT 25110 PCT | 88212-425635 including a multicopter plate and a plurality of multicopter rotors arranged on the multicopter plate, the plurality of multicopter rotors configured to generate thrust in a downward direction so as to provide lift to the aircraft. The plurality of multicopter rotors can be configured to be rotating in a VTOL mode of the aircraft and in a transition mode of the aircraft between VTOL and forward flight to provide lift to the aircraft, and can be configured to be shut off in a forward flight mode of the aircraft. The plurality of multicopter rotors can include four multicopter rotors. The multicopter plate can be arranged on the base frame assembly above the center of gravity of the aircraft. The multicopter plate can be arranged on the base frame assembly above the center of gravity of the aircraft such that a center of thrust of the plurality of multicopter rotors is arranged above the center of gravity.

[0033] An aircraft according to a further aspect of the present disclosure includes a base frame assembly, a multicopter assembly arranged on the base frame assembly, the multicopter assembly including a plurality of multicopter rotors configured to generate thrust in a downward direction so as to provide lift to the aircraft, and a main rotor assembly arranged above the multicopter assembly and including at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift. The multicopter plate can be arranged on the base frame assembly above the center of gravity of the aircraft.

[0034] In at least some embodiments, the multicopter assembly can include a multicopter plate, the plurality of multicopter rotors being arranged on the multicopter plate. The multicopter plate can be arranged on the base frame assembly above the center of gravity of the aircraft such that a center of thrust of the plurality of multicopter rotors is arranged above the center of gravity. The plurality of multicopter rotors can be configured to be rotating in a VTOL mode of the aircraft and in a transition mode of the aircraft between VTOL and forward flight to provide lift to the aircraft, and configured to be shut off in a forward flight mode of the aircraft. In some embodiments, the aircraft can further include a component adjustment assembly arranged on the base frame assembly and including a carriage configured to move relative to the base frame assembly. The main rotor assembly can be coupled to the carriage, and the carriage can be configured to be moved relative to the base frame assembly to move the main rotor assembly relative to the base frame assembly and thus move a center of pressure of the at least one wing relative to the center of gravity of the aircraft.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0035] A method according to a further aspect of the present disclosure includes providing a base frame assembly of an aircraft, arranging a multicopter assembly of the aircraft on the base frame assembly, the multicopter assembly including a plurality of multicopter rotors configured to generate thrust in a downward direction so as to provide lift to the aircraft, and arranging a main rotor assembly of the aircraft above the multicopter assembly, the main rotor assembly including at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift. The multicopter plate can be arranged on the base frame assembly below the center of gravity of the aircraft.

[0036] In some embodiments, the multicopter assembly can include a multicopter plate, the plurality of multicopter rotors being arranged on the multicopter plate. The plurality of multicopter rotors can be configured to be rotating in a VTOL mode of the aircraft and in a transition mode of the aircraft between VTOL and forward flight to provide lift to the aircraft, and configured to be shut off in a forward flight mode of the aircraft. In some embodiments, the method can further include arranging a component adjustment assembly of the aircraft on the base frame assembly, the component adjustment assembly including a carriage configured to move relative to the base frame assembly, and coupling the main rotor assembly to the carriage. The carriage can be configured to be moved relative to the base frame assembly to move the main rotor assembly relative to the base frame assembly and thus move a center of pressure of the at least one wing relative to the center of gravity of the aircraft.

[0037] A method of operating an aircraft according to a further aspect of the present disclosure includes operating an aircraft in a VTOL configuration in which at least one component of the aircraft is arranged relative to the base frame assembly such that a center of pressure of at least one wing of a main rotor assembly of the aircraft is approximately aligned with a center of gravity of the aircraft in a forward-aft direction, moving the at least one component relative to the base frame assembly to a forward flight configuration such that the center of pressure of the at least one wing is aft of the center of gravity of the aircraft, and operating the aircraft in a forward flight configuration.

[0038] In certain embodiments, the aircraft can include a base frame assembly, the main rotor assembly including the at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, the at leastAttorney Docket No.: MIT 25110 PCT | 88212-425635 one wing defining the center of pressure, and at least one component adjustment assembly arranged on the base frame assembly and configured to move at least one component of the aircraft relative to the base frame assembly.

[0039] In at least some embodiments, in the VTOL configuration, a leading edge of the at least one wing can face a direction of rotation of the main rotor assembly about the axis. The method can further include transitioning from the VTOL configuration of the aircraft to the forward flight configuration which includes rotating the at least one wing relative to an axis of rotation of the at least one wing such that the leading edge faces a forward direction of the aircraft for forward flight. The method can also further include, during the operating of the aircraft in the VTOL configuration, operating first and second counterbalance rotors of the aircraft that each face in the direction of rotation of the main rotor assembly. Even further, the method can include, during the transitioning from the VTOL configuration of the aircraft to the forward flight configuration, rotating one of the first and second counterbalance rotors to face a forward direction such that both of the first and second counterbalance rotors face the forward direction such that the first and second counterbalance rotors generate forward thrust for forward flight.

[0040] In some embodiments, the at least one component of the aircraft that is movable by the at least one component adjustment assembly can include the main rotor assembly. The method can further include moving the main rotor assembly via the at least one component adjustment assembly in a forward-aft direction to move the center of pressure of the at least one wing in the forward-aft direction. The method can further include moving a carriage of the at least one component adjustment assembly to which the main rotor assembly is coupled in the forward-aft direction relative to the base frame assembly such that movement of the carriage relative to the base frame assembly moves the center of pressure of the at least one wing forward and aft relative to the center of gravity of the aircraft. The method can also further include sliding the carriage along a linear rail of the at least one component adjustment assembly arranged on the base frame assembly.

[0041] In certain embodiments, the at least one component of the aircraft that is movable by the at least one component adjustment assembly can include a first component of the aircraft, and the method can further include moving the first component via the at least one component adjustment assembly in a forward-aft direction to move the center of gravity of aircraft in the forward-aft direction. The first component can be an electronics component ofAttorney Docket No.: MIT 25110 PCT | 88212-425635 the aircraft. The method can further include moving a carriage of the at least one component adjustment assembly to which the first component is coupled in the forward-aft direction relative to the base frame assembly such that movement of the carriage relative to the base frame assembly moves the center of gravity of the aircraft forward and aft relative to the center of pressure of the at least one wing.

[0042] In at least some such embodiments, the method can further include utilizing a state machine control scheme to operate the aircraft, the state machine control scheme including at least one flight state including at least one of a VTOL state, a forward flight state, at least one forward transition state for transitioning from VTOL to forward flight, or a at least one backward transition state for transitioning from forward flight to VTOL, and at least one operational safety state including at least one of a kill state, a disarm state, or an arm state. The control scheme can be configured to transition between the at least one flight state and the at least one operational safety state so as to transition the aircraft between VTOL flight and forward flight.

[0043] In some embodiments, the at least one forward transition state can include at least one of a rotor deceleration preparation state in which counterbalance rotors of the aircraft are temporarily disabled and reverse directions, a rotor deceleration state in which the main rotor assembly is brought to a stop, or a forward flight preparation state in which the at least one wing is rotated to a forward flight position and the main rotor assembly is moved to the aft position so as to position the center of pressure aft of the center of gravity, and the at least one backward transition state includes at least one of a quad hover state in which the at least one wing is rotated to a VTOL flight position and the main rotor assembly is moved to the forward position such that the center of pressure is approximately aligned with the center of gravity, or a rotor acceleration state in which the main rotor assembly is brought to a nominal operating speed.

[0044] In at least some embodiments, the control scheme can be further configured to at least one of transition to the kill state from the at least one flight state, from the disarm state, and from the arm state, transition to the disarm state from the kill state, or transition from the disarm state to the arm state. The control scheme can be further configured to at least one of transition from rotor deceleration state to the arm state or transition from the arm state to the quad hover state and the rotor spin up state. The control scheme can be further configured to at least one of transition from the quad hover state to the forward flight prep state and theAttorney Docket No.: MIT 25110 PCT | 88212-425635 rotor acceleration state, transition from the forward flight prep state to the forward flight state, transition from the rotor acceleration state to the VTOL state, transition from the rotor spin up state to the VTOL state, transition from the VTOL state to the deceleration prep state, transition from the deceleration prep state to the rotor deceleration state, transition from the rotor deceleration state to the forward flight prep state, or transition from the forward flight state to the forward flight prep state.

[0045] An aircraft according to a further aspect of the present disclosure includes a base frame assembly, a main rotor assembly including at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, the at least one wing defining a center of pressure, and at least one component adjustment assembly arranged on the base frame assembly and configured to one or more of move the center of pressure of the at least one wing relative to a center of gravity of the aircraft or move the center of gravity relative to the center of pressure.

[0046] In some embodiments, the at least one component adjustment assembly can be configured to move at least one component of the aircraft relative to the base frame assembly to one or more of move the center of pressure of the at least one wing relative to a center of gravity of the aircraft or move the center of gravity relative to the center of pressure. The at least one component of the aircraft that is movable by the at least one component adjustment assembly can include the main rotor assembly. The at least one component adjustment assembly can be a first component adjustment assembly and can be configured to move the main rotor assembly in a forward-aft direction to move the center of pressure of the at least one wing in the forward-aft direction. The first component adjustment assembly can include a first carriage configured to move in the forward-aft direction relative to the base frame assembly, and the main rotor assembly can be coupled to the first carriage of the first component adjustment assembly such that movement of the first carriage relative to the base frame assembly moves the center of pressure of the at least one wing forward and aft relative to the center of gravity of the aircraft. The first carriage can be movable between a forward position for VTOL flight of the aircraft in which the center of pressure of the at least one wing is approximately aligned with the center of gravity of the aircraft in the forward-aft direction and an aft position for forward flight of the aircraft in which the center of pressure of the at least one wing is aft of the center of gravity of the aircraft in the forward-aft direction. The first component adjustment assembly can further include a first linear railAttorney Docket No.: MIT 25110 PCT | 88212-425635 arranged on an upper side of the base frame assembly that extends in the forward-aft direction, and the first carriage can be slidably movable on the first linear rail.

[0047] In some such embodiments, the at least one component of the aircraft that is movable by the component adjustment assembly can further include a first component of the aircraft, the at least one component adjustment assembly further includes a second component adjustment assembly, and the second component adjustment assembly is configured to move the first component in a forward-aft direction to move the center of gravity of the aircraft in the forward-aft direction. The first component can be an electronics component of the aircraft. The second component adjustment assembly can include a second carriage configured to move in the forward-aft direction relative to the base frame assembly, and the first component can be coupled to the second carriage of the second component adjustment assembly such that movement of the second carriage relative to the base frame assembly moves the center of gravity forward and aft relative to the center of pressure of the at least one wing. The second carriage can be movable between an aft position for VTOL flight of the aircraft in which the center of pressure of the at least one wing is approximately aligned with the center of gravity of the aircraft in the forward-aft direction and a forward position for forward flight of the aircraft in which the center of pressure of the at least one wing is aft of the center of gravity of the aircraft in the forward-aft direction. The second component adjustment assembly can further includes a second linear rail arranged on an underside of the base frame assembly that extends in the forward-aft direction, and the second carriage can be slidably movable on the second linear rail. BRIEF DESCRIPTION OF THE DRAWINGS

[0048] The following Detailed Description references the accompanying drawings which form a part this application, and which show, by way of illustration, specific example implementations, in which:

[0049] FIG.1 is perspective view of a stop-rotor aircraft according a first aspect of the present disclosure, showing that the aircraft includes a base frame assembly, a main rotor assembly, a transition motor assembly, a motor assembly, a component adjustment assembly, a multicopter assembly, a counterbalance rotor assembly, and a tail assembly, the component adjustment assembly enabling movement of the main rotor assembly and the motor assemblyAttorney Docket No.: MIT 25110 PCT | 88212-425635 between forward and aft positions so as to move the center of pressure of the wings of the main rotor assembly relative to the center of gravity;

[0050] FIG.2 is a conceptual view of a conceptual aircraft and airfoils of the same in VTOL flight and forward flight;

[0051] FIG.3 is a conceptual view of a conceptual aircraft and airfoils of the same, showing the centers of gravity, pressure, thrust, and down force in VTOL flight and forward flight;

[0052] FIG.4 is a graphical view of an overview of performance of known VTOL aircraft in terms of hover productivity versus cruise productivity;

[0053] FIG.5 is a graphical view of an overview of performance of known VTOL aircraft in terms of hover efficiency as a function of disc loading;

[0054] FIG.6A is perspective view of a known helicopter aircraft;

[0055] FIG.6B is a perspective view of a known tiltrotor aircraft;

[0056] FIG.6C is a perspective view of a known tiltwing aircraft;

[0057] FIG.6D is a perspective view of a known tailsitter aircraft;

[0058] FIG.7A is a perspective view of a known stop-rotor aircraft, in particular a Herrick HV-2 Convertiplane;

[0059] FIG.7B is a perspective view of a known stop-rotor aircraft, in particular a Hughes Stopped Rotor;

[0060] FIG.7C is a perspective view of a known stop-rotor aircraft, in particular a Sikorsky X-Wing design;

[0061] FIG.7D is a perspective view of a known stop-rotor aircraft, in particular a Boeing / DARPA X-50A Dragonfly UAV;

[0062] FIG.8A is a perspective view of a known stop-rotor aircraft, in particular a Navy Stop-Rotor / Rotary Wing UAV;Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0063] FIG.8B is a perspective view of a known stop-rotor aircraft, in particular a Vargas- Clara and Redkar UAV;

[0064] FIG.8C is a perspective view of a known stop-rotor aircraft, in particular a SR800 by StopRotor Unmanned Aerial Systems;

[0065] FIG.8D is a perspective view of a known stop-rotor aircraft, in particular a JOBY LOTUS by JOBY Aviation;

[0066] FIG.9 is a conceptual view of forces that contribute to pitch torque on a general airplane, where the distances ℓPand ℓDdenote the distance between the center of gravity and the center of pressure of the wing and tail, respectively;

[0067] FIG.10 is a side view of the stop-rotor aircraft of FIG.1, showing the main rotor assembly in a VTOL position in which the leading edge of each wing is facing a direction of rotation of the main rotor assembly;

[0068] FIG.11 is a front view of the stop-rotor aircraft of FIG.10 in the VTOL position;

[0069] FIG.12 is a top view of the stop-rotor aircraft of FIG.10 in the VTOL position with the rotors of the counterbalance assembly each facing in opposing directions so as to provide stability during VTOL flight;

[0070] FIG.13 is a perspective view of the stop-rotor aircraft of FIG.10 in the VTOL position;

[0071] FIG.14 is a front perspective magnified view of the stop-rotor aircraft of FIG.10 in the VTOL position;

[0072] FIG.15 is a magnified top view of the stop-rotor aircraft of FIG.10 in the VTOL position not showing the tail assembly;

[0073] FIG.16 is a front view of the stop-rotor aircraft of FIG.15 in the VTOL position with the rotors of the counterbalance assembly each facing in opposing directions, showing the alignment of the centers of pressure of the wings, multicopter thrust, counterbalance thrust, and gravity along the axis of rotation;Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0074] FIG.17 is a side view of the stop-rotor aircraft of FIG.16 in the VTOL position with the rotors of the counterbalance assembly each facing in opposing directions, showing the alignment of the centers of pressure of the wings, multicopter thrust, counterbalance thrust, and gravity along the axis of rotation;

[0075] FIG.18 is a perspective view of the stop-rotor aircraft of FIG.1 in a forward flight position in which the second wing has been rotated such that the leading edge of the wing faces forward, and in which both rotors of the counterbalance assembly are facing forward, showing the assemblies prior to movement of the main rotor assembly and motor assembly rearward via the component adjustment assembly;

[0076] FIG.19 is side view of the stop-rotor aircraft of FIG.18 in the forward flight position;

[0077] FIG.20 is top view of the stop-rotor aircraft of FIG.18 in the forward flight position;

[0078] FIG.21 is a side view of the stop-rotor aircraft of FIG.18 in the forward flight position with the rotors of the counterbalance assembly each facing forward, showing the center of pressure of the wings being aft of the center of gravity, which is aft of the center of counterbalance thrust, and showing a center of pressure of the tail horizontal stabilizers;

[0079] FIG.22 is a perspective view of the components of a counterbalance subassembly of the counterbalance assembly of FIG.1, showing the rotor of the subassembly in a forward- facing position;

[0080] FIG.23 is a front perspective view of the base frame assembly and the multicopter assembly of the stop-rotor aircraft of FIG.1;

[0081] FIG.24 is a front perspective view of a main support frame of the base frame assembly, the motor assembly, and portions of the counterbalance assembly of the stop-rotor aircraft of FIG.1;

[0082] FIG.25 is an aft view of the main support frame of the base frame assembly, the motor assembly, and portions of the counterbalance assembly of the stop-rotor aircraft of FIG.1;Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0083] FIG.26 is a top view of the main support frame of the base frame assembly, the motor assembly, and portions of the counterbalance assembly of the stop-rotor aircraft of FIG.1;

[0084] FIG.27 is a top view of the main support frame of the base frame assembly, the component adjustment assembly, and portions of the counterbalance assembly of the stop- rotor aircraft of FIG.1;

[0085] FIG.28A is a side cross-sectional view of the main support frame and the component adjustment assembly of the stop-rotor aircraft of FIG.1 taken through line 28, 28B-28A, 28B in FIG.12, showing the main rotor assembly and the motor assembly in a forward position in which the center of pressure of the wings is aligned with the center of gravity;

[0086] FIG.28B is a side cross-sectional view of the main support frame and the component adjustment assembly of the stop-rotor aircraft of FIG.1 taken through line 28, 28B-28A, 28B in FIG.12, showing the main rotor assembly and the motor assembly in an aft position in which the center of pressure of the wings is aft of the center of gravity;

[0087] FIG.29A is a top view of a multicopter plate of the multicopter assembly of the stop-rotor aircraft of FIG.1;

[0088] FIG.29B is a perspective view of the multicopter plate of FIG.29A;

[0089] FIG.30 is a perspective view of the tail assembly of the stop-rotor aircraft of FIG. 1;

[0090] FIG.31 is a perspective view of the main rotor assembly of the stop-rotor aircraft of FIG.1, showing a central housing and the transition motor assembly arranged in the housing, and showing the second wing coupled to a servo of the transition motor assembly to enable rotation of the second wing from the VTOL position to a forward flight position;

[0091] FIG.32 is a top view of the main rotor assembly of FIG.31;

[0092] FIG.33 is a side perspective view of the main rotor assembly of FIG.31;

[0093] FIG.34 is a side view of a rotatable disc of the transition motor assembly of the main rotor assembly of FIG.31 with the second wing removed;Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0094] FIG.35 is a side view of a second wing base of the transition motor assembly of the main rotor assembly of FIG.31;

[0095] FIG.36 is a side view of a locking bumper and a locking pin of the transition motor assembly of the main rotor assembly of FIG.31, showing the locking bumper and the locking pin in the VTOL position of the second wing;

[0096] FIG.37 is an aft view of the locking bumper and the locking pin of the transition motor assembly of the main rotor assembly of FIG.31 in the VTOL position of the second wing;

[0097] FIG.38 is a side view of the locking bumper and the locking pin of the transition motor assembly of the main rotor assembly of FIG.31, showing the locking bumper and the locking pin in the forward flight position of the second wing, and showing the locking bumper engaging the locking pin to lock the rotational position of the main rotor assembly;

[0098] FIG.39 is an aft view of the locking bumper and the locking pin of the transition motor assembly of the main rotor assembly of FIG.31 in the forward flight position of the second wing;

[0099] FIG.40A is a perspective and top view of the multicopter plate of the multicopter assembly of the stop-rotor aircraft of FIG.1 with finite element analysis displacement results superimposed thereon;

[0100] FIG.40B is a perspective and top view of the multicopter plate of the multicopter assembly of the stop-rotor aircraft of FIG.1 with finite element analysis strain results superimposed thereon;

[0101] FIG.40C is a perspective and top view of the multicopter plate of the multicopter assembly of the stop-rotor aircraft of FIG.1 with finite element analysis stress results superimposed thereon;

[0102] FIG.41A is a front and top view of the motor assembly and the component adjustment assembly of the stop-rotor aircraft of FIG.1 with finite element analysis displacement results superimposed thereon;Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0103] FIG.41B is a front and top view of the motor assembly and the component adjustment assembly of the stop-rotor aircraft of FIG.1 with finite element analysis strain results superimposed thereon;

[0104] FIG.41C is a front and top view of the motor assembly and the component adjustment assembly of the stop-rotor aircraft of FIG.1 with finite element analysis stress results superimposed thereon;

[0105] FIG.42 is a computational fluid dynamics test result view of the stop-rotor aircraft of FIG.1;

[0106] FIG.43A is an overview of power distribution across the various electrical components in the stop-rotor aircraft of FIG.1;

[0107] FIG.43B is an overview of signal distribution across the various electrical components in the stop-rotor aircraft of FIG.1;

[0108] FIG.44 is a block diagram of a feedforward linearization method where r is the desired reference input, e is the error in current output and reference input, C(s) is the controller transfer function, d(s) are the disturbance dynamics, u(s) is the control input, Gp(s) is the plant transfer function, and y is the output of the system, and showing that the circles denote summing junctions where a (“−“) attached to the summing junction denotes a subtraction operator;

[0109] FIG.45 is a schematic overview of the developed Simscape multiphysics plant model used for computational simulations and controller validation for the stop-rotor aircraft of FIG.1;

[0110] FIG.46 is a block diagram of the disclosed control scheme for transforming desired state commands, xdes, into the appropriate control system to produce the set of controller outputs ξ which are then passed through a control allocation mixer to generate the set of motor commands;

[0111] FIG.47 is an overview of a body fixed coordinate frame and motor orientations of the stop-rotor aircraft of FIG.1;Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0112] FIG.48 is a further overview of the body fixed coordinate frame and motor orientations of the stop-rotor aircraft of FIG.1;

[0113] FIG.49 is a flow diagram of a Moore State Machine used for controlling the stop- rotor aircraft of FIG.1;

[0114] FIG.50 is a flow diagram overview of transition conditions from the operational states of the Moore State Machine of FIG.49 including a kill state;

[0115] FIG.51 is a flow diagram overview of transition conditions from the operational states of the Moore State Machine of FIG.49 including a disarmed state;

[0116] FIG.52 is a flow diagram overview of transition conditions from the operational states of the Moore State Machine of FIG.49 including an armed state;

[0117] FIG.53 is a flow diagram overview of transition conditions from the Quad Hover state of the Moore State Machine of FIG.49;

[0118] FIG.54 is a flow diagram overview of transition conditions from the VTOL states of the Moore State Machine of FIG.49 including a Rotor Spin Up;

[0119] FIG.55 is a flow diagram overview of transition conditions from the VTOL states of the Moore State Machine of FIG.49 including a VTOL;

[0120] FIG.56 is a flow diagram overview of transition conditions from the VTOL states of the Moore State Machine of FIG.49 including a Rotor Acceleration;

[0121] FIG.57 is a flow diagram overview of transition conditions from the transition states that take the control system from VTOL to forward flight of the Moore State Machine of FIG.49 including a Deceleration Preparation;

[0122] FIG.58 is a flow diagram overview of transition conditions from the transition states that take the control system from VTOL to forward flight of the Moore State Machine of FIG.49 including a Deceleration;

[0123] FIG.59 is a flow diagram overview of transition conditions from the transition states that take the control system from VTOL to forward flight of the Moore State Machine of FIG.49 including a Forward Flight Preparation;Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0124] FIG.60 is a flow diagram overview of transition conditions from the Forward Flight state of the Moore State Machine of FIG.49;

[0125] FIG.61 is a block diagram of a simulation pipeline for the Moore State Machine and control of the stop-rotor aircraft 10, where r is the desired reference input, e is the error, and y is the output of the system;

[0126] FIG.62 is a conceptual overview of various flight modes of the stop-rotor aircraft of FIG.1, showing the aircraft starting at rest, takeoff in VTOL and stable hover, transitioning to flip the second wing from VTOL to forward flight positions, reorienting the counterbalance rotors, moving the center of pressure of the wings via the component adjustment assembly, and forward flight;

[0127] FIG.63A is a graph of the stop-rotor aircraft response to varying commanded values in VTOL flight in a simulated environment, showing roll versus time;

[0128] FIG.63B is a graph of the stop-rotor aircraft response to varying commanded values in VTOL flight in a simulated environment, showing pitch versus time;

[0129] FIG.63C is a graph of the stop-rotor aircraft response to varying commanded values in VTOL flight in a simulated environment, showing yaw versus time;

[0130] FIG.63D is a graph of the stop-rotor aircraft response to varying commanded values in VTOL flight in a simulated environment, showing altitude versus time;

[0131] FIG.63E is a graph of the stop-rotor aircraft response to varying commanded values in VTOL flight in a simulated environment, showing position versus time;

[0132] FIG.63F is a graph of the stop-rotor aircraft response to varying commanded values in VTOL flight in a simulated environment, showing velocity versus time;

[0133] FIG.64A is a graph of the stop-rotor aircraft response to varying commanded values in forward flight in a simulated environment, showing pitch versus time;

[0134] FIG.64B is a graph of the stop-rotor aircraft response to varying commanded values in forward flight in a simulated environment, showing airspeed versus time;Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0135] FIG.64C is a graph of the stop-rotor aircraft response to varying commanded values in forward flight in a simulated environment, showing altitude versus time;

[0136] FIG.65 is a schematic overview of total deployed control system of the stop-rotor aircraft of FIG.1, showing a Pixhawk 1 running the pre-canned PX4 Standard VTOL controller and a Pixhawk 2 deploying a custom developed state machine controller of the Moore State Machine to run at 250 Hz;

[0137] FIG.66A is a schematic view of an RC Input Processing of a complete deployed Simulink controller for the stop-rotor aircraft 10 and control system of FIG.1;

[0138] FIG.66B is a schematic view of a Read, Write, and Process uORB Messages of the complete deployed Simulink of FIG.66A;

[0139] FIG.66C is a schematic view of the Moore State Machine of FIGS.49-60 and a Controller Output Post Processing of the complete deployed Simulink of FIG.66A;

[0140] FIG.66D is a schematic view of a PWM Write of the complete deployed Simulink of FIG.66A;

[0141] FIG.67A is a side view of a stop-rotor aircraft according a further aspect of the present disclosure, showing that the aircraft includes a base frame assembly, a main rotor assembly, a transition motor assembly, a motor assembly, a first component adjustment assembly, a multicopter assembly, a counterbalance rotor assembly, and a tail assembly, the first component adjustment assembly enabling movement of the main rotor assembly and the motor assembly between forward and aft positions so as to move the center of pressure of the wings of the main rotor assembly relative to the center of gravity, and showing that the aircraft includes a second component adjustment assembly configured to move a large component such as a battery pack forward and aft so as to move the center of gravity relative to the center of pressure;

[0142] FIG.67B is a side view of the stop-rotor aircraft of FIG.67A, showing the battery pack in a forward position after having been moved to this position via a linear actuator, carriage, and linear rail of the second component adjustment assembly;

[0143] FIG.68A is a bottom view of the stop-rotor aircraft of FIG.67A, showing the battery pack in an aft position;Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0144] FIG.68B is a bottom view of the stop-rotor aircraft of FIG.68A with the battery pack removed such that the linear actuator, an actuator housing, an actuator rod, the carriage, and the linear rail are visible;

[0145] FIG.69 is an aft view of the stop-rotor aircraft of FIG.67A, showing the base rail and central tongue of the linear rail and the groove of the carriage;

[0146] FIG.70 is a side view of the stop-rotor aircraft of FIG.67A in a forward flight position with the rotors of the counterbalance assembly each facing forward, showing the center of pressure of the wings being aft of the center of gravity after having been moved to this position via the first component adjustment assembly, which is aft of the center of counterbalance thrust, showing a center of pressure of the tail horizontal stabilizers, and showing the battery pack in a forward position such that the center of gravity is moved forward and further away from the center of pressure;

[0147] FIG.71 is a side view of an alternative arrangement of the stop-rotor aircraft of FIG.67A, showing that the aircraft without a first component adjustment assembly for the main rotor assembly and the motor assembly such that the main rotor assembly and the motor assembly are fixed in position relative to the base frame assembly of the aircraft; and

[0148] FIG.72 is a schematic diagram that shows a non-limiting example of a computing system that can be used to implement the techniques described herein. DETAILED DESCRIPTION

[0149] Certain embodiments will now be described to provide an overall understanding of the principles of the structure, function, manufacture, and use of the devices and methods disclosed herein. One or more examples of these embodiments are illustrated in the accompanying drawings. Those skilled in the art will understand that the devices and methods specifically described herein and illustrated in the accompanying drawings are non- limiting embodiments and that the scope of the present disclosure is defined solely by the claims. The features illustrated or described in connection with one embodiment may be combined with the features of other embodiments. Such modifications and variations are intended to be included within the scope of the present disclosure. A person skilled in the art will recognize how to rely upon the present disclosure to integrate, modify, and / or implement the embodiments of stop-rotor aircraft described herein and related methods of operating andAttorney Docket No.: MIT 25110 PCT | 88212-425635 controlling the same, into other configurations of aircraft, control systems and / or applications. To the extent features are described as being disposed on top of, below, next to, etc. such descriptions are typically provided for convenience of description, and a person skilled in the art will recognize that, unless stated or understood otherwise, other locations and positions are possible without departing from the spirit of the present disclosure.

[0150] Unless otherwise defined, all technical terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this disclosure belongs. Additionally, like-numbered components across embodiments generally have similar features unless otherwise stated or a person skilled in the art would appreciate differences based on the present disclosure and his / her knowledge. Accordingly, aspects and features of every embodiment may not be described with respect to each embodiment, but those aspects and features are applicable to the various embodiments unless statements or understandings are to the contrary.

[0151] Approximating language, as used herein throughout the specification and claims, may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “generally,” “about,” “approximately,” and “substantially” is not to be limited to the precise value specified. In some instances, the approximating language may correspond to the precision of an instrument for measuring the value. Here and throughout the specification and claims, range limitations may be combined and / or interchanged. Such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. In some non-limiting embodiments, such values modified by said terms may be in a range of plus or minus 5% of the indicated value.

[0152] ADDITIONAL BACKGROUND OF VTOL AIRCRAFT

[0153] The operational versatility offered by the ability of VTOL aircraft to takeoff and land from a stationary position has broadened aircraft applications to diverse environments with a broad range of mission requirements. VTOL aircraft can typically be classified into four main types: aircraft with a separate power plant for hover and forward flight, aircraft with a combined power plant for hover and forward flight, aircraft with an augmented power plant for hover, and aircraft with the same propulsion system for hover and forward flight.Attorney Docket No.: MIT 25110 PCT | 88212-425635 Aircraft with separate power plants for hover and forward flight use distinct engines optimized for each flight mode. This design allows for high efficiency and redundancy but requires carrying multiple power plants, which can significantly reduce payload capacity and increase operational costs. In contrast, aircraft with a combined power plant for hover and forward flight use a single engine for both functions. While this approach simplifies the propulsion system and improves weight efficiency, it often results in compromised performance because the engine is typically optimized for one mode over the other. Aircraft with augmented power plants for hover enhance their baseline propulsion capabilities with additional thrust sources, lift fans, or advanced technologies such as thrust vectoring. These augmentations improve hover efficiency and maneuverability but add extra weight and drag, which can reduce forward flight efficiency. Lastly, aircraft with the same propulsion system for both hover and forward flight utilize a single, versatile system designed to handle both flight modes. This design simplifies the aircraft and reduces weight but requires a highly efficient system that performs well in both vertical and horizontal thrust modes, often resulting in trade-offs in performance.

[0154] Generally, VTOL aircraft have increased mechanical and controller complexity compared to the purely hover or purely forward flight analogs. This increase in mechanical and / or controller complexity typically manifests as the aircraft performing more efficiently in hover or cruise, as shown in FIG.4. To deploy VTOL aircraft capable of long-range, fast, and high payload flight requires evaluation and optimization of efficiency across flight modes. The following provides definitions for efficiency across flight modes including hover and cruise.

[0155] Hover efficiency aims to capture how effectively an aircraft is using energy to hover. Typically, a suitable proxy for hover efficiency is disc loading, which is defined by Equation 1: Disc Loading ൌ Aircraft Weight ^kg^Rotor Area ^m2^ (1)Generally, lower disc loading indicates a weight or greater rotor area, thereby resulting in better lift generation and lower stall speeds. As shown in FIG.5, as disc loading decreases, hover efficiency increases. Conversely, higher disc loading implies a heavier aircraft with smaller wings, thereby resulting in faster cruise speeds and potentially greater structural efficiency.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0156] Cruise efficiency aims to capture how effectively aircraft generate lift in forward flight. The analogy of disc loading in the forward flight case is wing loading defined by Equation 2: Wing Loading ൌ Aircraft Weight ^kg^Wing Area ^m2^ (2)Wing loading typically correlates to stall speed, takeoff speed, and landing speed. For the purposes of the present disclosure, wing loading will qualitatively be used as a metric for cruise productively because from a design perspective, the size of the passive lifting surface relative to the weight is directly proportional to the cruise efficiency. Generally, lower wing loading suggests either a relatively light aircraft or relatively large wing area. Compared to similarly sized aircraft, those with a lower wing loading will require a slower cruise speed to maintain a stable altitude and are typically more maneuverable. Conversely, aircraft with a larger wing loading compared to an aircraft of a similar size will require faster cruising speeds. Therefore, wing loading will be used as a qualitative proxy for cruise efficiency for the purposes of evaluating VTOL aircraft performance in forward flight.

[0157] Some common VTOL aircraft that have a greater hover efficiency than cruise efficiency are described as follows. The helicopter, as shown in FIG.6A, is the flagship aircraft for achieving efficient hover: the large, thin, twisted blades are optimized for rotational performance. The high aspect ratio is optimized for rotational performance by maximizing the aerodynamic efficiency (i.e., lift / drag) and reducing drag on the rotor. The twist ensures that each section of the blade maintains an optimal angle of attack, thereby enhancing lift generation and overall rotor efficiency. However, regardless of the speed of the aircraft in forward flight, all of the lift must be provided by the continued spinning of the rotor. Furthermore, in forward flight, the retreating blade is subject to stall, while the advancing blade is subject to compressibility losses. In other words, the rotor uses more energy in forward flight to hold the aircraft at a stable altitude due to aerodynamic performance degradation and lack of a passive lifting surface. Evaluating Equation 1, the rotor area is relatively large given the long and thin blades, which would make disc loading low and thereby yield a high hover efficiency. However, evaluating Equation 2, the wing area is effectively zero, which would make the wing loading infinitely large and thereby infinitely inefficient in cruise.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0158] Some common VTOL aircraft that have a greater cruise efficiency than hover efficiency are described as follows. Tiltrotor aircraft typically have a passive forward flight lifting surface and distinct hover and lift configurations. With two rotors mounted on rotating engine nacelles, tiltrotors achieve vertical thrust in VTOL / hover by orienting the rotors directly upwards, as shown in FIG.6B. In forward flight, the engines rotate 90 degrees to generate thrust. Early tiltrotor aircraft were subject to significant rotor instabilities and can cause severe cockpit vibrations. Later tiltrotor aircraft were modified to address the vibration issues and the rotor instabilities, and to minimize operational loads on the cross shaft by mounting the engines and main transmission systems on the wingtips. Despite the tiltrotors success as a VTOL aircraft, tiltrotor concepts suffer from efficiency challenges due to using the same rotors across flight modes. Typically, tiltrotor engines are designed to lift the aircraft’s weight in VTOL, resulting in oversized engines for forward flight. Evaluating Equation 1, the rotor area is relatively small given the rotor area are approximately the size of fixed wing thrusters, which yields a larger disc loading and thereby less efficient in hover compared to vehicles with larger rotors (e.g., helicopter). However, evaluating Equation 2, the wing area is relatively large, which would make the wing loading relatively small and thereby relatively efficient in cruise compared to vehicles with small or no passive lifting wing. Therefore, tiltrotors are classified as cruise preferred due to the relatively high cruise efficiency and relatively low hover efficiency. In FIG.4, the tiltrotor falls to the right side of the desired region. This is due in practice to the engine being sized for VTOL and accepting the resulting engine inefficiency in forward flight thrust. For the purposes of classification, they will still be considered as cruise preferred due to the relatively lower wing loading compared to other VTOL.

[0159] A tiltwing aircraft is similar to a tiltrotor, except that rather than the rotors mounted on rotating nacelles, the entire wing is mounted on a pivot or hinge and rotates between VTOL and forward flight, as shown in FIG.6C. Early aircraft showed issues with propeller pitch control which caused aircraft to enter spins. Later aircraft stressed the importance of cross-shafting between engines to maintain redundancy and control in the event of an engine failure, as well as the necessity of direct propeller pitch control to ensure precise height and lateral control in VTOL. These takeaways informed the development of improved aircraft that successfully operated in conventional forward flight. However, the cross-link drive shaft introduced excessive vibrations, noise, and wing flexing. Similar to tiltrotors, tiltwingsAttorney Docket No.: MIT 25110 PCT | 88212-425635 typically have larger disc loading and smaller wing loading, classifying them as cruise preferred vehicles.

[0160] A tailsitter achieves VTOL by positioning the fuselage, wings, and tail of the aircraft vertically during takeoff, similar to a rocket or missile, as shown in FIG.6D. Once airborne, the aircraft transitions to forward flight by tilting the fuselage to the horizontal position, parallel to the ground. The pitching of the aircraft to transition between VTOL and forward flight is achieved by thrust vectoring with the engines, wings, or other control surfaces. Advancements in tiltrotor and other VTOL technologies led to tailsitter configurations falling out of favor, especially given stability and control issues through transition. These aircraft suffer the same engine mismatch as the tiltrotor and tiltwing aircraft where the engines are chosen to lift the vehicle in VTOL / hover, and are therefore oversized for forward flight. Also similar rational to tiltrotors and tiltwings, tailsitters typically have larger disc loading and smaller wing loading, classifying them as cruise preferred vehicles.

[0161] Stop-rotor aircraft have been explored for some time, some early designs of which are described as follows. The earliest design for a stop-rotor aircraft was the Herrick HV-2 Convertiplane. A biplane by construction, the upper wing of the Herrick was fixed for normal operation, but could be unlocked to allow the wing to auto rotate about the pylon and enable safe landing in an emergency, as shown in FIG.7A. The aircraft had issues with complexity and weight as compared to other VTOL methods. Another example of early designs includes the Hughes Stopped Rotor. The Hughes Stopped Rotor utilized a three- blade rotor with symmetric airfoils extending from each prong of the delta wing, as shown in FIG.7B. A delta wing was integrated into the rotor to provide passive lift in forward flight once the rotor stops. A further example of early designs include the four-bladed Sikorsky X- Wing design, shown in FIG.7C. This aircraft integrated a tip-powered rotor with four symmetric airfoil blades that utilized the Coandă effect to change the effective airfoil shape. Internal turbojets compressed air that passed to the tip of the rotor blade during hover and was exhausted out the back to provide cruise thrust. When sufficient speed is reached, the rotor would stop and lock into the X-Wing position.

[0162] An additional example of early stop-rotor designs includes the Boeing / DARPA X- 50A Dragonfly UAV, shown in FIG.7D. This aircraft utilized a single symmetric airfoil rotor system, also driven by turbine exhaust powered tip thrusters. Similar to previous stop-Attorney Docket No.: MIT 25110 PCT | 88212-425635 rotor aircraft constructions, the X-50A would divert exhaust through a tail nozzle to provide thrust for forward flight. The body was subject to unstable aerodynamic pitching moment due to airspeed and rotor wake that could not be stabilized by the flight controller.

[0163] An example of stop-rotor designs in smaller platforms includes a Navy Stop- Rotor / Rotary Wing UAV, shown in FIG.8A, that uses a two-bladed rotor driven by an electric motor that is stopped in forward flight once sufficient cruise speed is reached. Once the rotor is stopped, the retreating blade is rotated 180° along the blade spar to position the blade forward. The Navy Stop-Rotor aircraft provided the first prototype to successfully realize the potential of the stop-rotor aircraft utilizing the central lifting surface. However, the cruise efficiency in forward flight was limited with the usage of symmetric blades.

[0164] Another example of smaller stop-rotor designs includes the Vargas-Clara and Redkar, a stop-rotor UAV, as shown in FIG.8B. The design resembles that of a coaxial helicopter in that there are two rotors: a tail rotor and a wing rotor. The wing rotor utilizes a NACA 0012 where in helicopter mode, the wings rotate like a conventional rotor but only to supply torque to counter-act the motion of the tail rotor, and in forward flight the wing rotor fixes. The tail rotor is also a NACA 0012 and is the sole thrust generator in helicopter mode. In forward flight mode, the tail rotor acts as a conventional aircraft tail. At the front of the UAV is a pusher propeller that acts only to provide thrust in forward flight mode. The pusher propeller, tail rotor, and wing rotor are attached to a fuselage that houses the engine, rotor shaft, a clutch, and additional electrical and navigational hardware.

[0165] An additional example of smaller stop-rotor designs includes the SR800 design, manufactured by StopRotor Unmanned Aerial Systems, as shown in FIG.8C. Similar to the Boeing X-50A Dragonfly, the SR800 features a canard, as well as a large tail wing. Unlike the Boeing X- 50A Dragonfly, the rotor blade of the SR800 resembles that of a helicopter blade.

[0166] A further example of smaller stop-rotor designs includes the JOBY LOTUS, by JOBY Aviation, which features three rotors, two on the wing tips and one tilting propeller, as shown in FIG.8D. The wing tip propellers primarily provide lift in hover mode, but are also capable of tilting to initiate forward flight motion. As the aircraft begins to move forward, the tail mounted propeller begins to tilt to provide forward thrust. Once a suitable speed isAttorney Docket No.: MIT 25110 PCT | 88212-425635 reached, the wing tip rotors are stopped and reconfigured to extend the length of the fixed wing.

[0167] Finally, Brown and Ahuja present another example of a stop-rotor design, the AV-1 Hornet. Similar to the JOBY LOTUS, the AV-1 Hornet stops large wing-tip mounted rotors, rather than a single large spinning blade. While the wing-tip mounted rotors provide approximately 80% of the lift in hover, the design also consists of empennage mounted tilting ducted fans to augment the vertical lift in hover and provide thrust in forward flight. In forward flight, the wing tip mounted rotors are stowed such that the rotor blades function as active ailerons and extend the wingspan.

[0168] While the design of stop-rotor aircraft has seen advancements since the Herrick HV-2 was introduced, the modeling and control have been less developed but contributes equally to the long term success of stop-rotor aircraft. Vargas-Clara and Redkar derived a mathematical model for hover mode of a custom stop-rotor UAV platform and implemented linear control, optimal linear control, and nonlinear control in attempts of stabilizing the UAV in hover flight. To model their system, Newton’s second law is solved in vector form, and further expanded to Newton-Euler form. The linearized dynamics are shown in Equation 3: ^^0 1 0 0 0ì éെ^^0^^Ω^^^^^^^Ω^^ù ^^ ì ü (3)where yaw is independent of roll and pitch, but roll and pitch are coupled to one another. To control the linearized system, Vargas-Clara and Redkar attempted three different methods. First, a standard linear controller of the form shown in Equation 4 was attempted: x^ ൌ Ax ^ Bu^^^^, (4)where ẋ is the state rate, A is the state matrix, B is the controller matrix, and u(t) = Kx with K being the matrix of controller gains. Utilizing this method, a set of values for K such that theAttorney Docket No.: MIT 25110 PCT | 88212-425635 eigenvalues of [A − BK] have negative real parts could be found, suggesting a stable controller.

[0169] Second, optimal linear control was attempted, where the gains are chosen to minimize the following cost function shown in Equation 5: ^^൫^^^^^^,^^^^^^൯ ൌ^∙ x்^^^^^^ ∙ H ∙ x^^^^^^ ^ ^௧ ଶ ଶ^^x்^ ^^^^ ∙ ^^ ∙ ^^^^^^ ^ ^^ ∙ ^^ଶ^^^^^^^^,(5) suchmatrix, Q is the identity matrix scaled by 100, and R is the identity matrix. Using this method, gains that stabilize the system are provided.

[0170] Finally, a nonlinear controller to allow for stable control outside of equilibrium points was attempted. The Lyapunov direct method was utilized where a Lyapunov function V (x, t) was determined as shown in Equation 6: ^^^ ൌ െ2^^^^^ ଶ െ 2^^^^^ ଶ െ 2^^^^^ ଶ, (6)where a is any number lessrespectively. Therefore, Lyapunov stability for the system can be derived.

[0171] In summary, although the development of stop-rotor aircraft has undergone significant advancements from the Herrick HV-2, advancements in design, modeling, and control are necessary to bring the stop-rotor aircraft concept into everyday use. In the design, exploration into methods for handling aerodynamic mismatch and varying stability requirements between flight modes are still open areas. In modeling, generating refined models that capture the unique non-linear behaviors of stop-rotor aircraft are necessary to better understand the challenges associated with these vehicles. Finally, in controls, implementing, validating, and deploying control systems that are capable of stable transition on a stop-rotor aircraft are still desired. In unison, the development of design, modeling, and control are desired to achieve the successful forward and backward transition of a stop-rotor aircraft, all of which are addressed by the stop-rotor aircraft 10 of the present disclosure.

[0172] FIRST EMBODIMENT OF STOP-ROTOR AIRCRAFT

[0173] A stop-rotor aircraft 10 according to a first aspect of the present disclosure is shown in FIGS.1 and 10-43B. A control system 100 configured to control the stop-rotor aircraftAttorney Docket No.: MIT 25110 PCT | 88212-425635 described herein is shown in FIGS.44-66D. A stop-rotor aircraft 210, 210′ according to a further aspect of the present disclosure is shown in FIGS.67A-71.

[0174] The stop-rotor aircraft 10, also referred to herein as a stopped penta-rotor UAV or “SPERO,” is able to effectively transition between VTOL operation and directional flight operation, directional flight typically being forward flight, but also capable of being flight in other directions, such as backwards, sideways, and / or diagonal. Such directional flight is also able to occur separate from and / or while changing a vertical location with respect to the ground, and can be bi-directional during the VTOL to forward flight transition, as will be described in detail below. A number of different aspects of the disclosed designs of stop- rotor aircraft 10 enable the desired VTOL to directional flight transition, including the bi- directional VTOL to forward flight transition. These aspects include, for example, systems for changing a center of pressure 99A of the main wings 36, 38 such as the component adjustment assembly 60, 360, 360′, which can include, for example, a rotor adjustment assembly (i.e. a first component adjustment assembly) configured to move the main rotor assembly 30, 230 and the motor assembly 50, 250 so as to move the center of pressure 99A, 299A of the aircraft 10, 210, and a second component adjustment assembly (e.g., assembly 360, 360′ shown in FIGS.67A-71) configured to move a component relative to the base frame assembly 20, 220, 220′ so as to move the center of gravity 99C, 299C, 299C′ of the aircraft. The aspects can also comprise the inclusion of at least two propeller / wing assemblies—a main rotor assembly 30 and a multicopter assembly 70 (e.g., a quad-copter assembly) that is disposed below the main rotor assembly 30. In some embodiments, the main rotor assembly 30 can provide a majority of the lift for VTOL purposes, while the multicopter assembly 70 can stabilize, and provide some lift, during the transition between VTOL and directional flight.

[0175] The stop-rotor aircraft 10 is configured to achieve, at least, a certain set of operational capabilities, which include, for example, utilizing a central lifting surface that rotates in VTOL and fixes in forward flight, achieving stable operation in VTOL and forward flight, and achieving stable forward and backward transition between VTOL and forward flight. A person skilled in the art will understand that the term “VTOL” can encompass all aspects of vertically oriented flight modes, which can include vertical or partially vertical takeoff and landing, hovering, and similar modes. The term “forward flight” can encompassAttorney Docket No.: MIT 25110 PCT | 88212-425635 all aspects of forward oriented flight modes, which can include forward flight having pitch, roll, and yaw components.

[0176] Moreover, the stop-rotor aircraft 10 was designed, at least in the exemplary embodiment described herein, with a particular testing and usage environment in mind. For example, the stop-rotor aircraft 10 described herein was designed as a micro-UAV class for practical reasons: design, fabrication, and testing of a smaller drone is easier than a full-scale system. Micro-UAV class drones are defined as drones which weigh less than 2 kg, operate under 140 m altitude, and carry a payload of less than 1 kg. For other practical reasons, such as easy transportation, the drone size was limited to less than 500 mm in the largest dimension. For safety reasons while testing, the maximum speed of the drone was limited to 50 m / s. Finally, with the aforementioned specifications, parameters such as payload, flight time, and range were maximized compared to an equivalent drone in the micro-UAV class. Specifically, the maximum payload is greater than or equal to 1 kg, the maximum flight time is greater than or equal to 31 minutes, and the maximum range is greater than or equal to 50 km. These design parameters guided the design of the form of the UAV, as well as the avionics selection.

[0177] A person skilled in the art will appreciate that, although the description of the stop- rotor aircraft 10 focuses on the micro-UAV design, this design was utilized, at least in part, because of its effectiveness as a testing aircraft in controlled environments. A person skilled in the art will understand that the aspects of the design of the stop-rotor aircraft 10 as described herein is applicable to different types and configurations of aircraft, including, for example, larger-scale UAVs, both unmanned and manned, and aircraft of varying sizes including full-size passenger / cargo aircraft.

[0178] The design of the stop-rotor aircraft 10 was carried out in phases, with three versions of the aircraft being iterated on, leading to the final design described herein. A control co-design approach, defined as an integrated engineering system design method that modifies the physical design and control system to create new functionality and / or improve performance, was employed. The first design did not include a multicopter assembly nor a tail section, and instead only featured the characteristic central rotor and a servo rotated counterbalance system that provided counter torque in VTOL and thrust in forward flight. The aircraft of the first design lacked the passive pitch stability in forward flight that is typical of fixed wing aircraft. Furthermore, due to the geometry and number of the actuatorsAttorney Docket No.: MIT 25110 PCT | 88212-425635 in the first design, there was not adequate control authority for pitch and roll in forward flight.

[0179] Accordingly, the second design, which modified the first design, included an actuated center of pressure (i.e., movable, similar to the component adjustment assembly 60) to address pitch stability and control authority. Furthermore, a quadcopter architecture (i.e., similar to the multicopter assembly 70) was integrated to improve roll and thrust authority during the transition from VTOL to forward flight. The aircraft struggled to maintain acceptable stability, likely due to the multicopter assembly being far below the center of gravity of the aircraft. In the stop-rotor aircraft 10 described herein, the multicopter assembly 70 was raised on the aircraft 10 to be slightly above the center of gravity (e.g., center of gravity 99C in FIG.16). Raising the multicopter assembly 70 resulted in improved pitch and roll authority, as well as stability. The passive and dynamic stability of the aircraft 10 are covered in further detail below. The iteration on the design of the aircraft through the lens of controls enabled improvements in the necessary control authority and types of controllers to stabilize the vehicle.

[0180] Passive stability, defined as the tendency for a system to return to a stable state after a perturbation without additional control effort, is desirable as it improves the overall stability and simplicity of the necessary control system. In VTOL, the rotation of the top rotor has the potential to introduce highly nonlinear effects, especially if the axis of rotation is not aligned with the center of gravity and center of thrust of the counterbalance. As such, the configuration that aligns the center of gravity and center of thrust of the counterbalances on the axis of rotation, as shown in FIG.16 and FIG.17, is preferred. When aligned, the yaw balance equation for the body of the vehicle in VTOL is described by Equation 7: ∑^^௭ ൌ 0 ൌ ^^^^ െ ^^^^௧^^ ൈ ^^^ ^^௧^^, (7)where τCB is the torqueω^ rotor is the time rate of change of the angular speed of the rotor (i.e., the angular accelerationof the rotor). If τCB = Irotor × ω^ rotor, then the system is stable in yaw. The alignment of theaxis of rotation and counterbalance torque axis with the center of gravity ensures stability and robustness to disturbances. This concept is described in greater detail below.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0181] In forward flight for general aircraft, the center of lift is placed aft the center of gravity for passive pitch stability, such that when the nose of the aircraft pitches upwards, lift increases and provides a restoring moment, as shown in FIG.9. Conversely, when the nose of the aircraft pitches downwards, lift on the wing decreases and reduces the moment around the center of gravity. In addition to the relationship between the center of gravity and the center of pressure of the wing, the down force on the tail provides an additional stabilizing force: as the aircraft pitches nose up, the down force on the tail increases, providing a stabilizing moment. Furthermore, when the aircraft pitches nose down, the down force decreases, allowing the vehicle to stabilize back to trim. The configuration for a general fixed wing aircraft is shown in FIG.9. To produce a statically stable system, the Equation 8 should hold: ∑^^௬ ൌ 0 ൌ ^^^^^ௗ^௪^ െ ^^^^^^^^௧ , (8)where lDis the distancelPis the distance between the center of gravity and the center of pressure of the main wing, Fdown= 1 / 2ρcl,tailArefv2, and Flift= Fgravity. In the aircraft 10, the center of gravity, center of thrust, and center of pressure are not axially aligned due to the rotation of the wing, as can be seen in FIG.21. As such, the equation for a statically stable system in pitch is governed by Equation 9: ∑^^௬ ൌ 0 െ ℓ௬,்^^்^^௨^௧ െ ℓ௫,^^^^^^௧ ^ ℓ௬,^^^^^^^ ^ ℓ௫,^^^^^௪^, (9)where ℓy,T and thrust and center of pressure of the main wing, respectively, and ℓx,Pand ℓx,Dare the horizontal distances from the center of gravity to the center of pressure of the main wing and the tail, respectively. Furthermore, FThrustis the forward thrust force, FLiftis the lift force acting on the wing, FDrag is the drag force acting on the wing, and FDown is the down force acting on the tail.

[0182] The various components and systems of the stop-rotor aircraft 10 will now be described in detail with reference to FIGS.1 and 10-42. These components include, for example, a base frame assembly 20, a main rotor assembly 30, a transition motor assembly 40, a motor assembly 50, a component adjustment assembly 60, a multicopter assembly 70, a counterbalance rotor assembly 80, and a tail assembly 90. FIGS.1 and 10-21 show views ofAttorney Docket No.: MIT 25110 PCT | 88212-425635 the assembled stop-rotor aircraft 10 from various perspectives and in various configurations, including a configuration in which one of the wings 38 is in a VTOL position and the main rotor assembly 30 and the motor assembly 50 are at a forward position (FIGS.1 and 10-17), a configuration in which the same wing 38 is in a forward flight position and the main rotor assembly 30 and the motor assembly 50 are at a forward position (FIGS.18-20), and a configuration in which the same wing 38 is in a forward flight position and the main rotor assembly 30 and the motor assembly 50 are at an aft position (FIG.21). The various components that constitute the stop-aircraft 10, including the base frame assembly 20, the main rotor assembly 30, the transition motor assembly 40, the motor assembly 50, the component adjustment assembly 60, the multicopter assembly 70, the counterbalance rotor assembly 80, and the tail assembly 90, are shown in greater detail in FIGS.22-42.

[0183] In at least some embodiments, which include the micro-UAV design, at least some of these components may be formed by 3D printing with lightweight PLA (polylactic acid), although in other embodiments, the components can be 3D printed utilizing 3D printable plastic, metal, or carbon fiber. Other components, such as the servos and motors included in these assemblies, may be pre-fabricated.

[0184] In some embodiments, the wings 36, 38 may be formed of carbon fiber with PLA spars, while in other embodiments, may be formed of metal, foam, plastic, or other similar materials. In some embodiments, the various plate-like components (i.e., the support frame plates 23A, 23B, 23C, the base 33, the support trusses 35A, 35B, and the rotor plate 54) may be formed of carbon fiber, while in other embodiments, may be formed of metal, plastic, or other similar materials. In some embodiments, the counterbalance rods 84, 88 and the elongated rods 92 may be formed of carbon fiber, while in other embodiments, may be formed of metal, plastic, or other similar materials. In some embodiments, the horizontal stabilizers 94 may be formed of carbon fiber, while in other embodiments, may be formed of metal, plastic, or other similar materials. A person skilled in the art will understand that other similar materials and fabrication methods may be utilized for forming these components that would meet the design requirements of the specific aircraft.

[0185] In at least some embodiments, the weight of some of these assemblies may be broken down as follows: the base frame assembly 20 can weigh approximately 1.05 kg, the main rotor assembly can weight approximately 0.63 kg, the component adjustment assembly 60 can weigh approximately 0.35 kg, the counterbalance rotor assembly 80 can weighAttorney Docket No.: MIT 25110 PCT | 88212-425635 approximately 0.12 kg, and the tail assembly 90 can weight approximately 0.05 kg. A person skilled in the art will understand that these weights are merely representative of one exemplary embodiment of the stop-rotor aircraft 10, and any modifications in size and / or shape of these components that are derivable from the present disclosure would potentially affect the weight of the components.

[0186] The base frame assembly 20 of the aircraft 10 can be seen in detail in FIGS.1, 10, 11, 13, 14, and 23-27. As can be seen in the figures, the base frame assembly 20 includes a main support frame 22 and two legs 26, 27 extending downwardly therefrom, each of the legs 26, 27 being formed as an elongated tubular member and having a foot rod 28, 29 that provide a platform for the aircraft 10 to rest on such that the aircraft 10 can rest on a ground surface pre- and post-takeoff and landing. In some embodiments, the frame assembly 20 may include more or fewer legs depending on the design aspects of aircraft, such as more legs for embodiments having larger components. In some embodiments, the base frame assembly 20 may also include a landing gear (not shown) to further facilitate takeoff and landing.

[0187] As can be seen in greater detail in FIGS.24-27, the main support frame 22 can include three support frame plates 23A, 23B, 23C that are arranged in parallel and spaced apart from each other. The legs 26, 27 can be coupled to and extend downwardly from one of the plates, which can be, for example, the central, second support frame plate 23B. The first and second support frame plates 23A, 23B may be arranged spaced apart each other and define a space 24D therebetween. Various components may be arranged within the space 24, such as the servos 83, 87 of the two counterbalance rotor rods 84, 88 and portions of the component adjustment assembly 60. In some embodiments, the support frame plates 23A, 23B, 23C can be formed as thin carbon fiber plates, although other materials may be utilized based on desired design aspects and / or design requirements.

[0188] As shown in FIG.26 and FIG.27, the first and second support frame plates 23A, 23B can include four protrusions 24C that extend at 0°, 90°, 180°, and 270° from a center of the plate 23A, 23B. The first and second support frame plates 23A, 23B can each further include indentations 24B formed between the protrusions 24C. The indentations 24B can be formed complimentary to the contour of the inner support members 76A, 76B, 76C, 76D of the multicopter plate 72. The first and second support frame plates 23A, 23B can be fastened together as well as spaced apart at the forward and aft protrusions 24C via fasteners 25B, andAttorney Docket No.: MIT 25110 PCT | 88212-425635 can be fastened together as well as spaced apart at the side protrusions 24C via fasteners 25A.

[0189] As can be seen in FIGS.26-28B, the first support frame plate 23A includes an opening 24A formed therein. In some embodiments, the opening 24A is formed towards the aft protrusion 24C (i.e., the lower protrusion, as viewed in FIG.27) such that the linear actuator 62 and actuator rod 64 can be arranged at the forward protrusion 24C, as will be described in detail below. The opening 24A is located such that a carriage 66 of the component adjustment assembly 60 can extend from the actuator rod 64 and through the opening 24A. As a result, the carriage 66 along with the main rotor assembly 30 and motor assembly 50 that are attached thereto are able to move along the linear rail 67 which is arranged below the first support frame plate 23A and the opening 24A, as can be seen in FIGS.27-28B.

[0190] The third support frame plate 23C is arranged below the second support frame plate 23B, as can be seen in FIGS.10, 14, 19, 24, 25, 28A, 28B. Illustratively, the third support frame plate 23C has a smaller area than the first and second support frame plates 23A, 23B. The third support frame plate 23C can be coupled to the first and second support frame plates 23A, 23B at the forward and aft protrusions 24C via the fasteners 25B. In some embodiments, spacers 25C may be arranged around the fasteners 25B between the plates 23B, 23C. Illustratively, additional components of the aircraft 10, including onboard electronics such as a battery pack 96A and a flight computer / processor 96B, may be arranged on and / or supported by the third support frame plate 23C.

[0191] As can be seen in FIGS.1 and 10-20, and in greater detail in FIGS.31-35, the main rotor assembly 30 is the main rotating body of the aircraft 10 and houses the central lifting surface of the aircraft 10. The main rotor assembly 30 includes a central housing 32 and two wings 36, 38 coupled to and extending away from the central housing 32. Illustratively, one of the wings, a first wing 36, is fixed relative to the central housing 32, while the other wing, a second wing 38, can be rotated between a first position for VTOL flight (i.e., a “VTOL position” in which a leading edge 38A (FIG.12, 13, 19, 20) of the wing 38 faces in the direction of rotation of the rotor assembly 30) and a second position for forward flight (i.e., a “forward flight position” in which the leading edge 38A faces the same direction as the leading edge 36A of the first wing 36 in a direction of forward flight).Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0192] As can be seen in the side views of, for example, FIGS.10, 19, and 34-36, the wings 36, 38 include a symmetrical airfoil such that the second wing 38 includes the same airfoil properties (i.e., chord, maximum thickness, etc.) in both the VTOL and forward flight positions. Each wing includes a leading edge 36A, 38A and a trailing edge 36B, 38B. In the VTOL position, the leading edges 36A, 38A each face in the direction of rotation of the motor 52 (i.e., the leading edge 36A of the first wing 36A faces forward while the leading edge 38A of the second wing 38 faces aft, as viewed in the position shown in FIG.12). In the forward flight position, the leading edges 36A, 38A each face in the forward direction, as can be seen in FIG.20. In some embodiments, the wings 36, 38 may be fabricated via prepeg carbon fiber layup, although a person skilled in the art will understand that other types of wings that may be utilized with the aircraft 10, as well as other fabrication methods that can be utilized for manufacturing the wings.

[0193] Illustratively, the central housing 32 includes a base 33 formed as a plate, two support walls 34A, 34B extending upwardly away from the base 33, and two support trusses 35A, 35B, as shown in FIGS.31-35. Within the two support walls 34A, 34B and support trusses 35A, 35B, electronic devices for operation of the main rotor assembly 30 may be housed, which can include a communications module 35C for communicating with a base station, controller, or other communications and / or control systems configured to control the operation of the transition servo 41, as well as other components of the aircraft 10. A transition servo 41 or other type of motor for transitioning the second wing 38 from the VTOL position to the forward flight position and vice versa, which may be a component of the transition motor assembly 40, may also be arranged within the two support walls 34A, 34B and support trusses 35A, 35B and supported by the base 33.

[0194] As shown in FIGS.31-39, the transition motor assembly 40 includes the transition servo 41 and a rotatable disc 42 having a portion attached to the transition servo 41 and extending through the second support wall 34B. The rotatable disc 42 is driven by the transition servo 41. The rotatable disc 42 can include an axis of rotation that aligns with an axis of rotation of the wing 38. The rotatable disc 42 includes a plurality of holes 43 through which fasteners 43A may be inserted so as to couple a second wing base 48 to the rotatable disc 42. The second wing 38 is coupled to the second wing base 48 such that rotation of the rotatable disc 42 via the transition servo 41 causes rotation of the second wing 38, thus enabling transition of the second wing 38 from the VTOL position to the forward flightAttorney Docket No.: MIT 25110 PCT | 88212-425635 position and vice versa. The transition servo 41 should be capable of achieving a requisite angular position (i.e., angle 38C in FIG.21), which may be approximately 195° between a chord line 38D of the wing 38 and a forward-aft axis 38E of the aircraft 10. In some embodiments, this position may be held by the servo 41 at a high holding torque to keep the wing 38 actuated (i.e., held in the desired VTOL or forward flight position) during flight. A person skilled in the art will understand that the angle at which the wing 38 is held at should match the angle at which the first wing 36 is held, so the angle 38C can deviate from 195° in embodiments in which the first wing 36 is not held at 195°.

[0195] As can be seen in greater detail in FIGS.35-39, the transition motor assembly 40 further includes a locking bumper 44 arranged between the rotatable disc 42 and the second wing base 48. The locking bumper 44 may include a base portion 44A having a plurality of holes through which corresponding fasteners 43A can extend so as to fixedly secure the locking bumper 44 to the rotatable disc 42 and the second wing base 48. The locking bumper 44 may further include an outer portion 44B extending away from the base portion 44A, the outer portion 44B having a lower surface 45 configured to contact an upper platform 46A of a locking pin 46, as will be described below.

[0196] As can be seen in FIGS.36-39, the locking pin 46 includes the upper platform 46A and a terminal end 46B opposite the upper platform 46A. The locking pin 46 extends through an opening 33A formed in the base 33 of the central housing 32. A short hollow tube 47A having a circular upper flange 47B can be inserted into the opening 33A. The locking pin 46 can extend through the hollow tube 47A and can slidably move within the hollow tube 47A. The locking pin 46 can further include a spring 46C extending between and contacting a bottom surface of the upper platform 46A and an upper surface of the upper flange 47B of the hollow tube 47A, the spring 46C biasing the locking pin 46 into the position shown in FIG.36. The locking pin 46 can further include a stop pin 47C below the hollow tube 47A and configured to delimit upward movement of the locking pin 46 caused by the biasing of the spring 46C. In some embodiments, the locking pin 46 may be comprised of aluminum, although other materials may be utilized as would be understood by a person skilled in the art.

[0197] FIG.36 and FIG.37 show the locking bumper 44 in a first position corresponding to the VTOL position of the second wing 38, showing the locking bumper 44 with the second wing 38 and the second wing base 48 removed. The locking pin 46 is in a first positionAttorney Docket No.: MIT 25110 PCT | 88212-425635 within the base 33 of the central housing 32 in which the upper platform 46A is biased upwardly by the spring 46C a maximum distance above the base 33, as delimited by the stop pin 47C.

[0198] FIG.38 and FIG.39 shows the locking bumper 44 in a second position corresponding to the forward flight position of the second wing 38, having been moved to this position via the transition servo 41 having rotated the rotatable disc 42 in the clockwise direction as viewed in FIG.38. Similar to FIG.36, the locking bumper 44 is shown with the second wing 38 and the second wing base 48 removed. As the second wing 38 rotates from the VTOL position to the forward flight position, the lower surface 45 of the locking bumper 44 will eventually engage the upper platform 46A and push the locking pin 46 downwardly against the force of the spring 46C. As the second wing 38 approaches the forward flight position, the terminal end 46B of the locking pin 46 will enter a hole 54A (also see FIG.26) formed in the rotor plate 54 of the motor assembly 50, which will be described in greater detail below.

[0199] Because the terminal end 46B is arranged in the hole 54A in the second position, the locking pin 46 is fixed relative to the rotor plate 54, which is fixed in position relative to the axis 52A of rotation of the motor 52. Thus, the base 33 of the central housing 32 is prevented from rotating when the locking pin 46 is fixed relative to the rotor plate 54. As a result, the main rotor assembly 30, and thus the motor 52, can be temporarily locked in the forward flight position, thus preventing any undesired rotation of the second wing 38 about the axis 52A of rotation of the motor 52 during forward flight. The first wing 36 is similarly fixed in position relative to the central housing 32 in that the wing 36 is coupled to the housing 32 via a first wing base 49 that is fixedly coupled to the first support wall 34A of the central housing 32, as can be seen in FIG.31. A person skilled in the art will understand that, although the locking pin 46 is shown as a means of temporality locking the second wing 38 in place, other means known in the art may be utilized to lock the main rotor assembly 30 and the motor 52 in place during forward flight. Non-limiting examples of such means include magnetic components, high-friction components, a clutch device, a separate motor, and the like.

[0200] As can be seen in detail in FIGS.24-27, the motor assembly 50 includes the motor 52 and the rotor plate 54, as well as a bracket 55 for coupling the rotor plate 54 and the motor 52 to the carriage 66 of the component adjustment assembly 60. The motor 52 can be anyAttorney Docket No.: MIT 25110 PCT | 88212-425635 motor known to a person skilled in the art that is capable of high speed rotation of the main rotor assembly 30. A non-limiting example of a motor that could be utilized as the motor 52 is a Vertiq 40-06370 kV motor. This motor is advantageous in that it has the ability to switch between position and speed control. Notably, speed control is important for VTOL operation and position control is important for aligning the motor in forward flight. Furthermore, the 40-06 model was chosen because the torque specifications matched the expected torque necessary to spin the top rotor at the desired acceleration of 20 rad / s2.

[0201] Illustratively, the motor 52 can include a main rotating body 52B having an upper surface 52C, as shown in FIGS.24-27. The base 33 of the central housing 32 is coupled to the upper surface 52C via a plurality of fasteners. The motor 52 further includes a base 53 that is fixedly arranged on the rotor plate 54. The main rotating body 52B is configured to rotate relative to the base 53 so as to rotate the main rotor assembly 30. The motor 52 can further include communication means (not shown) configured to connect with a base station, controller, or other communications and / or control systems configured to control the operation of the motor 52.

[0202] Illustratively, as shown in FIGS.24-27, the bracket 55 can include a pair of support arms 56A, 56B coupled to and extending away from a bottom surface of the rotor plate 54, and a base platform 57 extending between and interconnecting the support arms 56A, 56B. The base platform 57 is fixedly coupled to an upper surface of the carriage 66 of the component adjustment assembly 60 such that the motor assembly 50 can be moved along the linear rail 67 via the carriage 66 sliding along the rail 67, as will be described below. In this way, the center of pressure 99A of the wings 36, 38 can be selectively adjusted via movement of the main rotor assembly 30 along the rail 67.

[0203] Illustratively, at least some components of the motor assembly 50 may be comprised of carbon fiber or similar materials that would be known to a person skilled in the art. For example, in some embodiments, the motor 52 may be pre-fabricated as described above. In some embodiments, the rotor plate 54 may be formed of carbon fiber, while in other embodiments, may be formed of metal, plastic, or other similar materials. In some embodiments, the bracket 55 may be 3D printed from a PLA material, while in other embodiments, may be formed of 3D printable metal, plastic, or other similar materials. A person skilled in the art will understand that other similar materials and fabrication methodsAttorney Docket No.: MIT 25110 PCT | 88212-425635 may be utilized for forming these components that would meet the design requirements of the specific aircraft.

[0204] As can be seen in FIGS.24-28B, the component adjustment assembly 60, which may also be referred to as a rotor adjustment assembly since the assembly 60 is capable of moving the main rotor assembly 30, includes a linear actuator 62, which may also be a linear servo. It is noted that the component adjustment assembly may also be referred to herein as a first component adjustment assembly. The component adjustment assembly 60 can further include a main actuator housing 63 and an actuator rod 64 extending outwardly of the main actuator housing 63 and movable relative to the main actuator housing 63. A linear servo may be beneficial in that it is lightweight while also having a high force output. A non- limiting example of a linear servo that may be utilized as the linear actuator 62 is a HLS12- 5050-6V. A person skilled in the art will understand that other means of actuating and moving the carriage 66 may be utilized, such as, for example, pneumatic, hydraulic, and similar actuators. Moreover, other means of moving the carriage 66 in the forward-aft direction would be understood by a person skilled in the art, such as, for example, rollers, belt drives, bearings, and the like.

[0205] The main actuator housing 63 is coupled to the first support frame plate 23A, in particular to the forward protrusion 24C. The actuator rod 64 extends outwardly from the main actuator housing 63 and is configured to be actuated by known means. For example, although not illustrated in detail, a person skilled in the art will understand that, in some embodiments, such an actuator housing 63 may include threads on an inner surface of an internal cavity that receives the actuator rod 64, and the actuator rod 64 has corresponding threads on an outer surface that engage the threads of the housing 63 so that rotation of the linear actuator 62 moves the actuator rod 64 toward and away from the housing 63.

[0206] As can be seen in greater detail in FIG.24 and FIG.25, the actuator rod 64 includes a terminal end 65 that is coupled to an upper surface of the base platform 57 of the bracket 55, which thus fixedly couples the carriage 66 to the actuator rod 64 via the connection of the bracket 55 to the carriage 66. The carriage 66 can include a groove 66A formed therein. The groove 66A opens downwardly and extends along a length of the carriage 66 relative to the forward-aft direction (i.e., the direction of travel of the actuator rod 64).Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0207] The component adjustment assembly 60 further includes the linear rail 67, as can be seen in FIG.24 and FIG.25. The linear rail 67 includes a base rail 67A and a central tongue 67B that both extend along the forward-aft direction. The base rail 67A is coupled to an upper surface of the second support frame plate 23B. A forward end of the linear rail 67 is located generally centrally on the second support frame plate 23B, as can be seen in FIG. 28A and FIG.28B. In some embodiments, the forward end of the linear rail 67 is located at a position such that the carriage 66 can be moved far enough forward to allow the center of pressure 99A of the wings 36, 38, the center of multicopter thrust 99B, the center of gravity 99BC, and the center of counterbalance thrust 99D can be aligned, as shown in FIG.17. This position of the carriage 66 is also shown in FIG.28A.

[0208] As can be seen in FIG.24 and FIG.25, the tongue 67B is received in the groove 66A of the carriage 66 so as to locate the carriage 66 on the linear rail 67 and enable slidable movement of the carriage 66 relative to the linear rail 67. In some embodiments, the tongue 67B can have a dovetail shape, and the groove 66A has corresponding recesses to receive the dovetail shaped tongue 67B, as shown in FIG.25. The carriage 66 may further include a lower surface 66B configured to slide along an upper surface 67C of the base rail 67A. In some embodiments, the base rail 67A may further include upwardly extending guide projections 67D that extend along a length of the base rail 67A on outer sides of the base rail 67A, as can be seen in FIGS.41A-41C. The opposing outer surfaces of the carriage 66 may slide along the inner surfaces of the guide projections 67D, providing additional stability during movement of the carriage 66 along the liner rail 67.

[0209] In operation, the linear actuator 62 actuates the actuator rod 64 so as to slidably move the carriage 66 along the linear rail 67. The carriage 66 is movable and lockable at any position along the rail 67 which thus allows for any distance between the center of gravity 99C and the center of pressure 99A to be achieved ranging from (i) the positions shown in FIG.17 and (ii) the aftmost position of the carriage 66 shown in FIG.21. In this way, because the main rotor assembly 30 and the motor assembly 50 are fixedly coupled to the carriage 66, the main rotor assembly 30 and the motor assembly 50 can be moved forward and aft relative to the base frame assembly 20, the multicopter assembly 70, the counterbalance rotor assembly 80, and the tail assembly 90. Illustratively, the carriage 66 along with the main rotor assembly 30 and the motor assembly 50, can be arranged in a forward position, as shown in FIG.28A. In some embodiments, in the forward position, asAttorney Docket No.: MIT 25110 PCT | 88212-425635 also illustrated in FIG.17, the center of pressure 99A of the wings 36, 38, the center of multicopter thrust 99B, the center of gravity 99BC, and the center of counterbalance thrust 99D are aligned at least in the forward-aft direction. The carriage 66, along with the main rotor assembly 30 and the motor assembly 50, can be moved rearward and thus be arranged in an aft position, as shown in FIG.28B. In some embodiments, in the aft position, as also illustrated in FIG.21, the center of pressure 99A of the wings 36, 38 is aft of the center of gravity 99C, which is aft of the center of multicopter thrust 99B. It is noted that the main rotor assembly 30 and the motor assembly 50 may be referred to as a “component” that is movable by the component adjustment assembly 60.

[0210] The capability to move the main rotor assembly 30 and the motor assembly 50 in this manner, which enables the adjustment of the center of pressure 99A of the wings 36, 38 forward and aft as shown in FIG.21, enables the pitch stability of the aircraft 10 to be actively modified in flight, thus allowing for precise stability adjustments during flight. As noted above, the principle that provides this stability is tied to the location of the center of pressure 99A aft being aft of the center of gravity 99C. Accordingly, so long as the center of pressure is located aft of the center of gravity 99C, stability is enhanced in forward flight. As will be described below, other configurations and methodologies of moving the center of pressure 99A aft of the center of gravity 99C are contemplated by the present disclosure.

[0211] FIG.21 also shows a vertical distance 99H (ℓy,T) between the center of counterbalance thrust 99D and the center of gravity 99C, a forward-aft distance 99G (ℓx,P) between the center of gravity 99C and the center of pressure 99A of the wings 36, 38, a vertical distance 99E (ℓy,P) between the center of gravity 99C and the center of pressure 99A, and a forward-aft distance 99F (ℓx,D) between the center of gravity 99C and the center of pressure 99B of the tail horizontal stabilizer 94. In some embodiments, the vertical distance 99E (ℓy,P), which is representative to the height of the main rotor assembly 30, may be approximately 55 mm to prevent interference with counterbalance rotors 85, 89. A person skilled in the art will understand that this distance may vary based on the various configurations of the other components of the aircraft 10 that might affect the center of gravity 99C and the center of pressure 99A.

[0212] Although the component adjustment assembly 60 is described herein as comprising linearly adjustable and movable components, a person skilled in the art will understand that other devices and methods of adjusting the center of pressure 99A of the wings 36, 38 are contemplated by the present disclosure. The main rotor assembly 30 and the motor assemblyAttorney Docket No.: MIT 25110 PCT | 88212-425635 50 could be arranged on any other component, alternative to or in addition to, the carriage 66 of the component adjustment assembly 60 that could enable movement of the center of pressure 99A, such as other linearly moving components or even rotational components. Moreover, the present disclosure contemplates movement of the center of pressure 99A in directions other than purely forward and aft, including side-to-side and upwardly and downwardly.

[0213] Similar to the motor assembly 50, the components of the component adjustment assembly 60 may be comprised of carbon fiber or similar materials that would be known to a person skilled in the art. In some exemplary embodiments, the linear rail 67 may be formed as separate components, in particular with the base rail 67A being formed by 3D printing with PLA material, or the other materials such as 3D printable carbon fiber, metal, or plastic, as described above, and the central tongue 67B can be pre-fabricated so as to match the carriage 66 which may also be pre-fabricated. The housing 63 and actuator rod 64 may also be pre-fabricated in some embodiments. A person skilled in the art will understand that other similar materials and fabrication methods may be utilized for forming these components that would meet the design requirements of the specific aircraft.

[0214] Because the loads of the wings 36, 38 will pass through the motor assembly 50 and the component adjustment assembly 60, the structure of the motor assembly 50 and the component adjustment assembly 60 was designed and analyzed using finite element analysis (FEA) to ensure adequate strength. The results of the FEA are shown in FIGS.41A-41C. The maximum stress across the carbon fiber structures of the motor and linear rail assemblies 50, 60 is approximately 20 MPa, which is well below the yield strength of carbon fiber. The maximum stress in the 3D printed bracket 55, including the support arms 56A, 56B and the base platform 57, is approximately 5 MPa, which is within a safety factor of the yield strength. Therefore, the construction of the bracket was assumed to be sufficient to transmit the expected loads.

[0215] As shown in FIGS.1, 10-20, 23, 29A, and 29B, the multicopter assembly 70, also referred to herein as a quadcopter in embodiments having four rotors, includes a multicopter plate 72 having four rotor coupling portions 73A, 73B, 73C, 73D located at the four corners of the plate 72, and one multicopter rotor 74A, 74B, 74C, 74D arranged at each coupling portion 73A, 73B, 73C, 73D. A person skilled in the art will understand that, although the multicopter plate 72 is described herein as including four multicopter rotors 74A, 74B, 74C,Attorney Docket No.: MIT 25110 PCT | 88212-425635 74D, other numbers of multicopter rotors, such as two, three, five, six, seven, eight, or any other number of rotors capable of producing the same or similar lift effects on the aircraft 10 as the multicopter assembly 70 described herein. Although the multicopter rotors 74A, 74B, 74C, 74D are illustrated as two-bladed rotors, other numbers of blades on the rotors 74A, 74B, 74C, 74D may be utilized as would be understood by a person skilled in the art based on the aircraft design requirements, such as, for example, rotors having three blades, rotors having four blades, or rotors having more than four blades. It is noted that the computational models and simulations described below utilize properties related to three-bladed rotors. The principles evidenced by the results of these models and simulations applies to the disclosed aircraft regardless of the number of blades on the rotors.

[0216] As can be seen in FIG.29A and FIG.29B, the multicopter plate 72 includes four inner support members 76A, 76B, 76C, 76D that extend between outer coupling portions 78A, 78B, 78C, 78D of the plate 72 so as to define an inner open space 79 of the plate 72. As described above, inner support members 76A, 76B, 76C, 76D can be formed complimentary to the contour of the indentations 24B of the first and second support frame plates 23A, 23B. In some embodiments, each of the outer coupling portions 78A, 78B, 78C, 78D is configured to couple to accommodate various components of the stop-rotor aircraft 10. For example, the outer coupling portion 78A may be a forward coupling portion 78A with holes for the fasteners 25B and a hole for coupling to at least a portion of the linear actuator 62. The outer coupling portion 78C may be an aft coupling portion 78C with holes for the fasteners 25B. The outer coupling portions 78B, 78D may be a side coupling portion 78B, 78D with holes for the fasteners 25A as well as each including a side hole through which a corresponding counterbalance rotor rod 84, 88 can extend.

[0217] Illustratively, as shown in FIG.29A and FIG.29B, the multicopter plate 72 can further include a plurality of outwardly extending support members 77A, 77B, 77C, 77D, where a group of support members 77A, 77B, 77C, 77D extends outwardly from a corresponding inner support member 76A, 76B, 76C, 76D to a corresponding one of the four rotor coupling portions 73A, 73B, 73C, 73D (i.e., a group of support members 77A extends outwardly from the inner support member 76A to the outer coupling portion 73A). In some embodiments, the support members 77A, 77B, 77C, 77D can converge as they extend from the corresponding inner support member 76A, 76B, 76C, 76D to the corresponding rotor coupling portion 73A, 73B, 73C, 73D so as to form a triangular shape when viewed in theAttorney Docket No.: MIT 25110 PCT | 88212-425635 top-down direction, this shape increasing the strength and rigidity of the rotor coupling portion 73A, 73B, 73C, 73D for supporting the multicopter rotors 74A, 74B, 74C, 74D thereon.

[0218] As can be seen in FIGS.1, 10-20, and 23, a multicopter rotor 74A, 74B, 74C, 74D is coupled to each rotor coupling portion 73A, 73B, 73C, 73D. Illustratively, each multicopter rotor 74A, 74B, 74C, 74D is driven by a unique motor. In some embodiments, the motors may be brushless motors, although other motors for driving the rotors 85, 89 could be utilized as would be understood by a person skilled in the art. A non-limiting example of a motor configured to be utilized with the multicopter rotors 74A, 74B, 74C, 74D is a MAD 24071900 kV motor with a T-Motor F35A electronic speed controller (ESC). For the quadcopter configuration, the motors should be able to accommodate a 4 to 1 thrust to weight ratio. The maximum takeoff weight for the system is constrained to be approximately 2 kg. The maximum thrust of the chosen quadcopter motors is approximately 2.1 kg with bladed 5055 propellers, therefore providing an estimated thrust to weight ratio of 4:1. Furthermore, the maximum current rating for the motors is 49.4 A. The chosen ESCs are rated for 35A of continuous current and 45A peak current.

[0219] Because the multicopter assembly 70 is utilized at least to provide an additional lift component to the aircraft 10, each of the multicopter rotors 74A, 74B, 74C, 74D is oriented in a downward-facing direction such that rotation of the rotors 74A, 74B, 74C, 74D generates thrust in the upward direction. In some embodiments, the multicopter rotors 74A, 74B, 74C, 74D can be reversible such that the multicopter rotors 74A, 74B, 74C, 74D can be rotated in both rotational directions, one that generates upward thrust and one that generates downward thrust.

[0220] As can be seen in FIGS 10-17, the multicopter plate 72 is arranged between the first and second support frame plates 23A, 23B and is coupled thereto via the fasteners 25A, 25B. This position of the multicopter plate 72 and the associated multicopter rotors 74A, 74B, 74C, 74D on the aircraft 10 causes the center of multicopter thrust 99D to be slightly above the center of gravity 99C, as can be seen in FIG.16 and FIG.17. As described above, the center of multicopter thrust 99D being above the center of gravity 99C results in improved pitch and roll authority, as well as stability.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0221] As can be seen in top-down views of FIGS.10, 15, and 20, the multicopter assembly 70 is arranged centrally relative to the main rotor assembly 30 and the motor assembly 50 when the main rotor assembly 30 and the motor assembly 50, and thus the carriage 66, are arranged in the forward position along the linear rail 67. In this way, the center of multicopter thrust 99D can be aligned along the axis of rotation of the main rotor assembly 30 and the motor assembly 50, as shown in FIG.16 and FIG.17. When the main rotor assembly 30 and the motor assembly 50 transition to the aft position during forward flight, sufficient lift is created by the wings 36, 38. In this position, the multicopter rotors 74A, 74B, 74C, 74D can be shut off as the additional lift provided by the rotors 74A, 74B, 74C, 74D may no longer be desired. This is why no center of multicopter thrust 99D is shown in FIG.21.

[0222] In some embodiments, the multicopter assembly 70 can further include communication means (not shown) configured to connect with a base station, controller, or other communications and / or control systems configured to control the operation of the multicopter rotors 74A, 74B, 74C, 74D. In some embodiments, the multicopter rotors 74A, 74B, 74C, 74D may be controlled in unison via such a control system or may be individually controllable for precise thrust adjustments.

[0223] As described above, the multicopter assembly 70, and in particular the multicopter rotors 74A, 74B, 74C, 74D, can at least stabilize the aircraft 10, provide lift during transitions (i.e., transitions from VTOL flight to forward flight and vice versa), and can add redundancy. The multicopter assembly 70 not only plays an important role in maintaining stability throughout flight operations, which will be described in detail below, but also decreases the overall disk loading of the system compared to an equivalent system without the multicopter assembly 70. As the various components of the multicopter assembly 70, such as the multicopter rotors 74A, 74B, 74C, 74D, should be relatively spaced apart, the design of the multicopter plate 72 was chosen to balance weight and displacement on the plate 72 under loading. FEA was performed on the plate 72 to determine loading under maximum multicopter operation, the results of which are shown in FIGS.40A-40C.

[0224] In addition to supporting the multicopter rotors 74A, 74B, 74C, 74D, the multicopter plate 72 can also serve as an electronics bay dedicated to housing many of the onboard electronics (not shown in detail) of the aircraft 10. For example, certain electrical components, including certain sensors, radios, flight computers, and power distributionAttorney Docket No.: MIT 25110 PCT | 88212-425635 systems, may be arranged in the space 79. Some of these electronics can also be housed on the third support frame plate 23C as opposed to within the space 79.

[0225] As can be seen in FIGS.1, 10-27, including greater detail in FIG.22, the counterbalance rotor assembly 80 includes a first counterbalance rotor subassembly 82 and a second counterbalance rotor subassembly 86 located on an opposing side of the aircraft 10. Each of the first and second counterbalance rotor subassemblies 80, 86 includes a servo 83, 87 that can be mounted between the first and second support frame plates 23A, 23B via servo mounting brackets 83A, 87A. Each servo 83, 87 includes a counterbalance rotor rod 84, 88 extending therefrom, and each servo 83, 87 is configured to rotate the rod 84, 88 about its central lengthwise axis. A non-limiting example of a servo 83, 87 that can be utilized to rotate the rods 84, 88 is a AGFRC B26CLM. This servo has a compact form factor that allows it to fit within the first and second support frame plates 23A, 23B.

[0226] Illustratively, a first terminal end of the counterbalance rotor rod 84, 88 is coupled to the servo 83, 87 and a second terminal end of the rod 84, 88 is coupled to a counterbalance rotor 85, 89. In some embodiments, the counterbalance rods 84, 88 may have a length of 200 mm to ensure adequate spacing between the rotors 85, 89 and the base frame assembly 20, the main rotor assembly 30, and the multicopter assembly 70, and to ensure adequate yaw authority. In some embodiments, the counterbalance rods 84, 88 may have a length in a range of 150 mm to 250 mm, and in some embodiments, may have a length in a range of 100 mm to 300 mm. A person skilled in the art will understand that the length of the counterbalance rods 84, 88 may depend on the dimensions of the remainder of the aircraft 10, and thus may vary based on design parameters and other properties of the aircraft 10.

[0227] Although the multicopter rotors 74A, 74B, 74C, 74D are illustrated as two-bladed rotors, other numbers of blades on the rotors 74A, 74B, 74C, 74D may be utilized as would be understood by a person skilled in the art based on the aircraft design requirements, such as, for example, rotors having three blades, rotors having four blades, or rotors having more than four blades. It is noted that the computational models and simulations described below utilize properties related to three-bladed rotors. The principles evidenced by the results of these models and simulations applies to the disclosed aircraft regardless of the number of blades on the rotors.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0228] The counterbalance rotor rods 84, 88 can each extend through a corresponding hole in one of the side outer coupling portions 78B, 78D, and may be supported therein via a bushing (not shown). The counterbalance rotor rods 84, 88 may be formed of carbon fiber in some embodiments, although other materials are contemplated by the present disclosure as would be understood by a person skilled in the art. Each counterbalance rotor 85, 89 includes a unique motor 85A, 89A that is fixedly coupled to the counterbalance rotor rod 84, 88. In some embodiments, the motors may be brushless motors, although other motors for driving the rotors 85, 89 could be utilized as would be understood by a person skilled in the art. A non-limiting example of a motor 85A, 89A that can be utilized with the rotors 85, 89 is the MAD 24071900 kV motor with the T-Motor F35A ESC, as described above.

[0229] Rotation of the counterbalance rotor rod 84, 88 will rotate the counterbalance rotors 85, 89 about the central lengthwise axis of the rotor rod 84, 88. In this way, the counterbalance rotors 85, 89 can be rotated between and including a forward-facing position (e.g., FIGS.1, 10-12, and 18) and an aft-facing position (e.g., FIGS.13, 15, and 16, showing the counterbalance rotor 89 is the aft-facing position). In other embodiments contemplated by the present disclosure, the counterbalance rotor rod 84, 88 may be fixed and the motors 85A, 89A and rotors 85, 89 themselves may be configured to rotate about the counterbalance rotor rod 84, 88 between and including the forward-facing and aft-facing positions. Illustratively, the second counterbalance rotor 89 may face in the aft direction in a VTOL configuration in which each counterbalance rotor 85, 89 faces the direction of rotation of the main rotor assembly 30.

[0230] In operation, the counterbalance rotors 85, 89 serve a dual purpose. In VTOL, which can include hovering, the counterbalance rotors 85, 89 provide yaw torque to compensate for the effects of rotation of the main rotor assembly 30 and improve yaw authority. During forward flight, the counterbalance rotors 85, 89 can be aligned in the forward-facing position and thus provide forward thrust.

[0231] As can be seen in FIGS.1, 10, 12, and in greater detail in FIG.30, the tail assembly 90 includes elongated rods 92 that are coupled to and extend away from the main support frame 22, in particular from the aft protrusion 24C of the first and second support frame plates 23A, 23B in some embodiments. The assembly 90 further includes a stabilizer hub 93 and horizontal stabilizers 94 extending away from the hub 93. As can be seen in FIG.21, the horizontal stabilizers 94 provide stabilizing downward force in forward flight (i.e., center ofAttorney Docket No.: MIT 25110 PCT | 88212-425635 pressure 99B of the horizontal stabilizers 94). It is noted that the tail assembly 90 placement depends on the statics of the system under stable flight, as well as the wake of the other components such as main support frame 22, the main rotor assembly 30, the motor assembly 50, the multicopter assembly 70, and the counterbalance rotor assembly 80, also referred to as the “body” of the aircraft 10.

[0232] Computational fluid dynamics (CFD) simulation of the body of the aircraft 10 was performed to estimate the distance of the wake, which was determined to be approximately 165 mm from the back of the body in the illustrative embodiment, as shown in FIG.42. Therefore, the statics problem solving for the tail assembly 90 location was lower bounded at 165 mm. Beyond ensuring that the tail assembly 90 is out of the wake of the body of the aircraft 10, evaluating Equation 9 subject to the chosen tail assembly 90 parameters yields ltail= 400mm.

[0233] As described at various points above, in operation, the stop-rotor aircraft 10 is configured to be operated in various flight configurations and operating modes. For example, in one flight configuration, the second wing 38 is in the VTOL position and the main rotor assembly 30 and the motor assembly 50 are at the forward position along the linear rail 67, as shown in FIGS.1 and 10-17. In this flight configuration, which may be referred to as a first flight configuration, the center of pressure 99A of the wings 36, 38, the center of multicopter thrust 99B, the center of gravity 99BC, and the center of counterbalance thrust 99D are approximately aligned along the axis of rotation, as can be seen in FIG.16 and FIG.17. In this flight configuration, both counterbalance rotors 85, 89 are facing the direction of rotation about the axis of rotation (e.g., FIGS.13, 15, and 16) for VTOL flight. This may also be referred to as a VTOL flight configuration. It is noted that any configuration in which the center of pressure 99A is approximately aligned along the axis of rotation with the center of gravity 99C may be referred to as a VTOL flight configuration in at least some embodiments.

[0234] In another flight configuration, which may be referred to as a second flight configuration, the second wing 38 is in a forward flight position and the main rotor assembly 30 and the motor assembly 50 are at the forward position along the linear rail 67, as shown in FIGS.18-20. In this flight configuration, both counterbalance rotors 85, 89 are facing the forward direction. This may be a flight configuration of the aircraft 10 during a transition between VTOL and forward flight.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0235] In a further flight configuration, also referred to as a third flight configuration, the second wing 38 is in a forward flight position (e.g., FIGS.18-20), the main rotor assembly 30 and the motor assembly 50 are at an aft position (FIG.21). In this flight configuration, the multicopter rotors 74A, 74B, 74C, 74D are shut off, and the center of pressure 99A of the wings 36, 38 is aft of the center of gravity 99BC, which is aft of the center of counterbalance thrust 99D. In this flight configuration, both counterbalance rotors 85, 89 are facing the forward direction for forward flight. In some embodiments, the multicopter rotors 74A, 74B, 74C, 74D may remain on and rotating in order to provide additional lift. This flight configuration may also be referred to as a forward flight configuration. It is noted that any configuration in which the center of pressure 99A is aft of the center of gravity 99C may be referred to as a forward flight configuration in at least some embodiments.

[0236] As noted above, the aircraft 10 can include various onboard electronics, which can include avionics that manage and control various functions of the aircraft 10. The avionics can include a power distribution system and method as well as a signal distribution system and method. It is noted that, in the illustrative embodiment, the electronics arranged on components other than the main rotor assembly 30 completely separate from and decoupled from the electronics arranged in the main rotor assembly 30 due to the complexity of transmitting power through a motor rotating at high speeds. With the development of ultra high speed slip rings and inductive power transfer, the rotor and body electronics could eventually be unified in some embodiments.

[0237] The power distribution system provides reliable operation and delivery of power to the onboard electronics, and ensures that each component is receiving the appropriate voltage and current necessary for nominal operation. An exemplary configuration of the power distribution electronics utilized in the illustrative embodiment of the stop-rotor aircraft 10 is summarized in FIG.43A, with the electronics for the “main body system” depicted below the dashed line (i.e., electronics for any component not including the main rotor assembly 30, which includes electronics for the motor assembly 50, the component adjustment assembly 60, the multicopter assembly 70, and the counterbalance rotor assembly 80) and electronics for the main rotor assembly 30 depicted above the dashed line. A person skilled in the art will understand that other configurations of electronics and the signal distributions of the same, including those depicted in FIG.43A and 43B, may be utilized depending on the aircraft and control system designs, and as the available avionic systems evolve over time.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0238] On the main body system, the core of the power distribution system is a Power Distribution Board (PDB), which takes in power from a 6S Lipo battery nominally at 22.2 V and routes it to the various onboard modules. Specifically, the PDB supplies 22.2 V to all the electronic speed controllers to power the brushless motors (i.e., the motor 52 and the motors for the rotors 74A, 74B, 74C, 74D, 85, 89). The PDB also supplies 22.2 V to a 7.4 V step down converter, which provides a stable voltage to the onboard actuators and servos (i.e., the linear actuator 62, the servos 83, 87). The PDB also powers a Pixhawk Power Module, which is a 5 V switching regulator, which provides a stable power source for a Pixhawk 6X flight controller. The flight controller directly powers an RC receiver and telemetry radio. Finally, the PDB supplies 22.2 V to a 5 V step down converter that powers a Raspberry Pi. The Raspberry Pi powers a second Pixhawk flight controller through a USB cable. Two Pixhawk modules 75 may be arranged in the base frame assembly 20, as shown in FIG.23.

[0239] On the main rotor assembly 30, a 2S Lipo battery, which may be arranged in the central housing 32, nominally provides 7.4 V directly to the servo 41 that actuates the second wing 38, as well as a 5 V regulator. The 5 V regulator is a Buck Converter that powers a Seeed RP2040 microcontroller which in turn powers the RC receiver.

[0240] The signal distribution system is responsible for routing the control and data signals. The core of the signal distribution system are the Pixhawk flight controllers. The controller system and the integration of the developed controller with the greater flight controls stack will be described in greater detail below. The signal distribution scheme is summarized in FIG.43B, with the main body system depicted below the dashed line and the main rotor assembly 30 depicted above the dashed line.

[0241] To initiate action on the aircraft 10, the user sends a command via a handheld transmitter on the ground station. The transmitter sends a 2.4 GHz radio frequency signal across 16 channels, which is received by an X8R receiver on the main body system and by an X8R receive on the main rotor assembly 30. The main body system receives signals on channels 1 through 8, which is then sent via SBUS to the first Pixhawk flight controller. To control the various motors onboard, the Pixhawk flight controllers writes a pulse width modulated (PWM) signal at a 500-2500 μs interval for the counterbalance servos 83, 87 and a 1000-2000 μs interval for all other motors (i.e., the motor 52, the linear actuator 62, and the motors for the rotors 74A, 74B, 74C, 74D, 85, 89). The ESCs convert the PWM signal into desired motor speed using the built in microcontrollers, which then generate a high-Attorney Docket No.: MIT 25110 PCT | 88212-425635 frequency series of voltage pulse signals that approximate sinusoidal waveforms. The pulses control switching of metal-oxide semiconductor field effect transistors within the ESC, which creates three separate waveforms 120 degrees out of phase from one another that are directed to the motors (i.e., motors for the rotors 74A, 74B, 74C, 74D, 85, 89). As noted above, the ESCs can be a T-Motor F35A ESC is some embodiments.

[0242] On the main rotor assembly 30, the second X8R receiver receives the 2.4GHz radio frequency signal from channels 9 - 16 of the transmitter on the ground station. The X8R receiver interprets the 2.4GHz signal into a 1000-2000 μs PWM signal dictated by historic standardization. However, to fully rotate the second wing 38, the servo 41 should achieve 270° actuation. This actuation range is achievable with a 500-2500 μs PWM signal. Therefore, the onboard Seeed microcontroller reads the 1000-2000 μs PWM signal and expands the range to 500-2500 μs to provide the control signal to the servo 41.

[0243] For unrestrained flight tests, local position can be provided through a Vicon Motion capture system. Motion capture data can be packaged and sent to the Raspberry Pi using ROS 1 over WiFi. The Raspberry Pi can transform the motion capture data from the local motion capture frame of East, North, Up (ENU) to Front, Right, Down (FRD), which is preferred by the PX4 flight control software. Once the data is transformed into the preferred reference frame, the Raspberry Pi publishes the data over serial using MAVROS for the Pixhawk controller to read. The Pixhawk controller then fuses the position and orientation data into the onboard extended Kalman filter to update the estimated position and orientation of the aircraft 10.

[0244] The specific electronics hardware described above may include the following additional subcomponents. For example, built into the Pixhawk 6X can be three ICM-45686 inertial measurement unit (IMU) sensors with vibration rejection and temperature stability for triple redundancy, two barometers (ICP20100 and BMP388), and one BMM150 magnetometer. These three sensing units are used to generate position and orientation estimates.

[0245] Telemetry radios can be used to transmit flight information, such as position, speed, altitude, and system status, between the Pixhawk flight controller and the ground control station. Holybro 100 mW 915MHz SiK telemetry radios can be used for their compatibility to the Pixhawk 6X. A RX / TX pair can be used for manual piloting by sending desiredAttorney Docket No.: MIT 25110 PCT | 88212-425635 setpoints (e.g., altitude, speed, etc.) to the control system of the aircraft 10. An important factor in choosing a radio pair is the ability to communicate with two receivers to allow for isolated electronics between the main rotor assembly 30 and the main body system. To this effect, a Taranis X9D transmitter and a FrSky X8R receiver can be chosen for RC communication. Finally, a Raspberry Pi 4 can be used to relay motion capture data from a Vicon system to the flight controller. The Raspberry Pi 4 with 8 GB of SDRAM cab be used for its relatively low cost and versatile operation.

[0246] CONTROL SYSTEM FOR STOP-ROTOR AIRCRAFT

[0247] A control system configured to control the stop-rotor aircraft 10, in particular to at least control and stabilize the aircraft 10 in forward and backward transitions between VTOL and forward flight, is described below. First, the derived analytical models of the system are described, namely models for the yaw and altitude dynamics as they are identified to be heavily influenced by nonlinear effects such as the quadratic dependence of rotor speed on lift and drag. Next, the development and implementation of the computational Simulink model is described for use in controller development. Finally, a complete control system for stabilizing the aircraft 10 across transitions is disclosed, including evaluation and tuning in simulation.

[0248] Models of the system, even low fidelity ones, help provide an understanding of the expected dynamics and provide insight into the control strategies necessary to achieve the desired behaviors. In derivation of the control system, both analytical and computational models were developed. Specifically, analytical models were developed to provide mathematical insight and closed form stability margins for the system. Computational models were used to capture a broader range of behaviors to test, tune, and evaluate control systems on a systems scale. Each model type is reviewed in further detail below.

[0249] Analytical models were used to derive closed-form mathematical insights into the stability of key behaviors of the system. Primarily, analytical models were derived for the yaw on the main body system subject to the rotation of the main rotor assembly 30 in VTOL and the altitude of the aircraft 10 subject to change in speed of the main rotor assembly 30 or the quadcopter 70. Although mostly analyzing simplified forms of the system, each of the models indicated some form of nonlinearity and / or instability.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0250] Yaw dynamics were the first area of interest for deriving an analytical model because the high inertia and acceleration / deceleration of the main rotor assembly 30 were identified as factors heavily impacting the stability of the aircraft 10. By solving Newton’s second law of motion for the yaw of the main body system in VTOL, the following equation for the angular acceleration of the main body system is derived as Equation 10: ^^^ ^^ௗ௬^^^^^ௗ௬^^^^ ൌ ^^^^௧^^ ∙ ^^^^^௧^^^^^^ ^ଶ^^^^ௗ^^^^^^^ଷ^^^^௧^^^^^^ଶ ^ ^^^^^^^^, (10)where Ibodyacceleration of the main body system, Irotor is the moment of inertia of the main rotorassembly 30, ^^^ rotor (t) is the angular acceleration of the main rotor assembly 30, ρ is thedensity of air, cd is the coefficient of drag of the top wings 36, 38, r is the distance from the axis of rotation to the center of pressure of the wing, ωrotor(t) is the angular velocity of the main rotor assembly 30, and τCB (t) is the torque acting on the main body system due to the counterbalances 85, 89.

[0251] To evaluate system stability, typically a Laplace transform would be performed to generate a transfer function of the system. However, the Laplace transform of Equation 10 cannot be explicitly computed unless a form for ωrotor (t) is assumed. For the purposes of what is feasible to operate the aircraft 10, two cases are explored: first, constant main rotor assembly 30 speed and second, constant main rotor assembly 30 acceleration.

[0252] In the first case for constant main rotor assembly 30, in stable hover, if it is assumedthat ωrotor (t) = ωrotor is a constant and ^^^ rotor (t) = 0, then Equation 10 can be reduced toEquation 11 as follows: ^^^ ^^ௗ௬^^^^^ௗ௬^^^^ ൌଶ ^^^^ௗ^^^^^^^ଷ^^ଷ^^௧^^ ^ ^^^^^^^^, (11)where ωrotor no longer produces the following Equation 12: ^^ ^^^^ ൌ^^^ೌ^ ^ ఛ^ಳ^^^^^ௗ௬(12)where s is the Laplace ωbodyyaw in the Laplace domain and cdrag = 1 / 2ρcd Aref r3ωrotor2.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0253] To generate a closed loop transfer function, some form for τCBshould be assumed. In the simplest form, τCB = u(s), where u(s) is defined by the feedforward linearization diagram depicted in FIG.44. If the disturbance is taken to be Equation 13: ^^^^^^ ൌ െ^^^ೌ^^ , (13) which is simply d(t) = −cdrag in theFIG.44 is described in Equation 14: ^^^^^ ^^^^^^ ൌ^ ^^^^ାூಳ^^^ೞ, (14) where C(s) is an arbitrary(s) is determined by the poles of the system. Namely, s = − C(s) / Ibody. Because Ibody will always be > 0 for physical systems, so long as C(s) is designed to be stable, the closed loop system will be stable.

[0254] In the second case of constant rotor acceleration, while the stable hover condition provides useful insights into the in-flight stability, the plant will start at rest and must transition between rest and rotation during aircraft 10 transitions during flight. As such, certain operations will be important where the main rotor assembly 30 speed is not constantand the main rotor assembly 30 is accelerating. If it is assumed that ^^^ rotor (t) = caccel and ωrotor= c ^ t, then Equation 10 can be reduced to Equation 15:accel ^^ ^^^ ^^^^ ൌ ^^ ⋅ ^^ ^^ ^^ௗ௬ ^^ௗ௬ ^^௧^^ ^^^^^ଶ^^^^ௗ^^^^^^^ଷ^^^^^^^^ ⋅ ^^^ଶ ^ ^^^^^^^^, (15)whichூ^^^^⋅^ ^^^^^ఛ^ೞ^^ ೌ^^^^ ^ಳ^ ^ ^^ ^^ ൌ^ ^ , (16)^^ௗ௬మ రூ ^ ூ ^ ூ3where c = ρc A r .comp d ref accel

[0255] Similar to constant main rotor assembly 30 speed in the first case, some form for τ should be assumed to generate a closed loop transfer function. If τ = u(s), where u(s) isCB CB defined by the feedforward linearization diagram depicted in FIG.44, and the disturbance is taken to be Equation 17: మ ூ ^ ^ ା^^^^^^ ೌ^^^^ ^^^^^ ^ ^^ ^^, (17) ^Attorney Docket No.: MIT 25110 PCT | 88212-425635 then the closed loop transfer function reduces to the form in Equation 14 and exhibits the same stability criteria as the first case. In the time domain, d(t) = −(Irotorcaccel + ccomp t2 / 2).

[0256] Altitude dynamics were the second area of interest for generating analytical models due to the nonlinear relationship between main rotor speed and lift. By solving Newton’s second law of motion for the vertical acceleration of the main body system in VTOL, the following equation for acceleration in z of the main body system is derived in Equation 18: ^^ ⋅ ^^^^^^^ ൌ ^^ ⋅ ^^ െ ^ ଶ ଶଶ^^^^^^^^^^^^ ^^^^௧^^^^^^ െ ^^^௨^ௗ^^^^, (18)where z(t) is the altitude of the aircraft 10, m is the mass of the main body system, g is the gravitational constant, cl is the coefficient of lift of the wings 36, 38, and Fquad(t) is the vertical force contributed by the quadcopter 70. Similar to the yaw dynamics, unless a form of ωrotor (t) is assumed, the Laplace transform for Equation 18 cannot be explicitly calculated or usefully manipulated. As in yaw dynamics, two forms of ωrotor(t) can be assumed: first, constant main rotor assembly 30 speed and second, constant main rotor assembly 30 acceleration.

[0257] In the first case for constant main rotor assembly 30 speed, in a stable hover, if ω (t) is assumed to be constant such that ω (t) = ω , then Equation 18 is simplified torotor rotor rotor Equation 19: ^ଶଶ^ ^ ^ ^ ^^ ⋅ ^^^^^ ൌ ^^ ⋅ ^^ െ^^^^ ^^ ^^ ^^െ ^^ ^^ , (19)^ ^^^ ^^௧^^ ^௨^ௗଶand produces the ி^^^^ ^ೠೌ^ೡ^^^^ ^ ^^ ^^ ൌെ , (20)య మ^^ ^^3wherein c = 1 / 2ρc A r .drag d ref rotor

[0258] Similar to yaw dynamics, a closed loop transfer function for the system can be derived by assuming some form of F (s). Similar to yaw dynamics, if F (s) = u(s),quad quadwhere u(s) is defined as shown in the feedforward linearization diagram (FIG.44), then the disturbance can be taken to be Equation 21: ^ೡ^^^^ ^ ^^ ^^ ൌ, (21) ^Attorney Docket No.: MIT 25110 PCT | 88212-425635 which translates to d(t) = cvertin the time domain. The closed-loop transfer function can then be defined as Equation 22: ^^௬^^^^^^^^^^^^ ൌ ^^^^^ൌ^^^^ି^^మ, (22) where C(s) is the controllerthe additional s term in the denominator compared to the closed-loop yaw dynamics, the controller should be chosen to stabilize the additional integrator.

[0259] In the second case of constant main rotor assembly 30 acceleration, as the aircraft 10 should accelerate and decelerate the main rotor assembly 30 to complete transitions, the first case as described above will not always hold. However, if the angular acceleration of the main rotor assembly 30 is assumed to be constant such that ωrotor = caccel ^ t, then Equation 18 is simplified to Equation 23: ^^ ∙ ^^^^^^ ൌ ^^ ∙ ^^ െ ^ ଶଶ^^^^^^^ ଶ^^^^^ ^^^^^^^^ ∙ ^^^ െ ^^^௨^ௗ^^^^, (23)which produces the^^^^^^ ൌ ^^ య െ^^^^,భி^ೠೌ^^^^ ^ ^^రെ ^^మ(24) where ccomp= ρcdArefr3caccel2.

[0260] Similar to the first case, if Fquad(s) = u(s), where u(s) is defined as shown in the feedforward linearization diagram FIG.44, and the disturbance is defined as Equation 25: ^^^^^^ ൌ^^^మି^^^^^,^^య(25)then the closed loop transfer domain, d(t) = mg − ccomp,lt2 / 2. Similar to the constant speed main rotor assembly 30 case, the stability of the closed loop transfer function will depend on choosing C(s) to stabilize the extra integrator in Equation 22.

[0261] In summary, the analytical models provided insight into operational guidelines for the main rotor assembly 30, as well as stability criteria. For both yaw dynamics and altitudeAttorney Docket No.: MIT 25110 PCT | 88212-425635 dynamics, when the main rotor assembly 30 speed or main rotor assembly 30 acceleration are set to be constant, stability criteria can be defined for the closed-loop system.

[0262] While analytical models described above provide specific details about stability conditions, they may lack information about the performance of the entire system. Furthermore, deriving closed form equations for the entire system is rather costly and complicated. Instead, computational models were used to test the closed loop stability of the developed controller and perform initial tuning. A computational model for the aircraft 10 was derived using the Simscape Multibody package, which forms joints between 3D bodies to formulate and solve the equations of motion in the time domain. The shape and properties of the 3D bodies of the aircraft 10 have been translated from CAD models utilizing the Inventor plugin. The overall derived plant model is provided in FIG.45.

[0263] In addition to the default imported model and generic world physics, several components were added and modified to improve the fidelity of the model including aerodynamic forces, ground contacts, and frame transformations to imitate the data that will be written from PX4 to the Simulink controller. Each of these added physics will be covered in further detail below. Body Parameter Value Notes ρ 1.293 kb / m3Standard temperature and pressure Cl 0.75 NACA 0012 at 5° angle of attack Main Wing Cd0.025 NACA 0012 at 5° angle of attack Aref 0.056 m2Wing planform area Cl1 Tail lift coefficient Tail Wing Cd 0.1 Tail drag coefficient Aref0.010 m2Tail planform area Cd 1 Coefficient of drag of a plate Main Body Aref0.030 m2Body frontal area TABLE 1

[0264] Aerodynamic forces include lift acting on the main wings 36, 38 and the tail assembly 90, drag acting on the main wings 36, 38, tail assembly 90, and main body system, and propeller dynamics. Lift acting on the main wings 36, 38 varied depending on the aircraft 10 operating mode (i.e., VTOL vs forward flight). In forward flight, the standard liftAttorney Docket No.: MIT 25110 PCT | 88212-425635 equation was used for lift on the main wings 36, 38 and the tail assembly 90, as shown in Equation 26: ^^^^^^^^ ൌ ^ଶ ^^^^^^^^^^^^ଶ, (26) where ρ is the density of air, clis the lift coefficient, Arefis the reference area, and v is the velocity of the main body system relative to the air. It was assumed that the aircraft 10 was operating without wind disturbances, so the velocity was taken to be the velocity of the vehicle in the main body system fixed x direction. Similarly, the standard drag equation was used to compute the drag on the main wing 36, 38 and tail assembly 90, as shown in Equation 27: ^^^^^^^^ ൌ ^ଶ ^^^^ௗ^^^^^^^ଶ(27) where cdis the drag coefficient. The aerodynamics parameters used to populate Equations 26 and 27 are summarized in Table 1.

[0265] In VTOL, the lift on the wings 36, 38 was calculated using the equations for a rotating tapered wing provided by Equation 28: రయ^^^^^^^^ ൌ ^^^^^ ଶ^^^^^^ି^^ర^^^^ ^ ^^^^^^ି^^య^^^^ ^^^^^௧^^^^^൬ ^ ൬ ^^, (28)where ω (t)change in chord, c, over the length of the wing, l, defined such that c(l) = ml + b. For the carbon fiber wings 36, 38, m = −0.384 and b = 0.16. Furthermore, rtipis the distance between the axis of rotation and the tip of the wing 36, 38, and rroot is the distance between the axis of rotation and the root of the wing 36, 38. Similarly, the drag on the wing 36, 38 was inferred from the lift equation of a rotating wing shown in Equation 29: రయ^^^^^^^^ ^^^^ ି^ ^^^^ ^^ଶ^^^ ^ర^^^^ ^^^^^^ି^^య^^^^ ^^^^^ ^.Equations 26-29 were were then applied to the Simscape bodies using the External Force and Torque block.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0266] Finally, the propellers (i.e., rotors 74A, 74B, 74C, 74D, 85, 89) dynamics were handled with the Rotor block from the Simulink Aerospace blockset. The Rotor block takes in speed and outputs force and moment from the associated propeller using a quadratic function of speed, as shown in Equation 30: ^^ ଶ௫,௬,௭ ൌ ^0,0,െ^^்ఘ^^ ^^ଶ൧^^ ൌ 0,0,െ^ ଶ ଶ௫,௬,௭ ^ ^ொఘ^^ ^^ ൧, (30)where CTis the thrustpropeller, R is the propeller radius, and ω is the propeller rotational speed, defined by the input to the block. As noted above, all of the propellers / rotors modeled in the stop-rotor model are three-bladed 5055 propellers with a CT = 0.0107, CQ = 7.83 × 10−4, and R = 63.5 mm. A person skilled in the art will understand that two-bladed (as figures),three-bladed, four-bladed, or rotors with even more blades may be on aircraft 10 and modeled based on the equations and methodology described herein.

[0267] Ground contacts were integrating to capture takeoff and landing behaviors from a stationary position. The ground plane was created using the Infinite Plane block connected to an exported convex hull of the stop-rotor aircraft 10 main body system from the associated Simscape solid body through a Spatial Contact Force block. The Spatial Contact Force block provides normal and friction forces between the plane and the convex hull.

[0268] Once the control system is deployed on the real system (i.e., the physical aircraft 10), it will pull position and orientation estimates from the Pixhawk extended Kalman filter, which fuses data from three accelerometers / gyroscope (ICM-20649, ICM-42688-P, ICM- 42670-P), one magnetometer (BMM150), two barometers (BMP388), and vision estimation from the external motion capture system in unrestrained flight tests. A person skilled in the art will understand that other numbers and specific models of these components, such as accelerometers, gyroscopes, magnetometers, barometers, and motion capture systems may be utilized based on design requirements of the aircraft 10.

[0269] It is important to ensure that the position and orientation estimates used to tune the controllers match the reference frame of the flight controller. The quaternion, velocity, and position data from the Pixhawk extended Kalman filter are reported in the north east down (NED) frame. The quaternion is defined as the rotation from the NED initial earth frame toAttorney Docket No.: MIT 25110 PCT | 88212-425635 the main body system fixed frame. Velocity and position data are reported with respect to the inertial earth frame.

[0270] A summary of the disclosed high level control architecture is depicted in FIG.46. In essence, a desired set point xdes is passed into a controller 110, which then produces a set of desired forces and torques, ξ. These desired forces and torques are then passed into a control allocation mixer 120 to map the forces and torques into motor commands Θ. This control scheme was chosento the nature of the different states of the aircraft 10. Across most states, the base controller 110 can remain the same, but the mixer 120 must change due to the significant structural changes of the plane (e.g., counterbalance rotors 85, 89 rotating direction 180°). The implementation and structure of each of these modules is covered in further detail below.

[0271] For all VTOL operation and transitions, a simple Proportional-Integral-Derivative (PID) Controller can be used as the controller 110 due to the overall simplicity and ease of tuning and implementation. Feedforward linearization modifications were also included to address the dynamics derived in the analytical models for the yaw and altitude dynamics and their related two cases described above. For the forward flight operation, the coupling between pitch and change in lift makes the simultaneous control of airspeed and height through standard PID challenging.

[0272] A total energy control system (TECS) can be used as the base controller 110 for forward flight states to handle the coupling between airspeed, height, and pitch in a simple and proven effective manner. In a TECS, the controller integrates the various aspects of the aircraft that contribute to kinetic and potential energy (e.g., airspeed, altitude, thrust) to formulate a multi-input multi-output system to control the aircraft in a stable and decoupled manner. In essence, TECS utilizes throttle to control the total energy of the system and pitch to control the distribution between kinetic and potential energy. In utilizing a TECS, set points are transformed into the energy domain.

[0273] The mixer provides a linear transformation between the controller 110 outputs and the desired speeds and positions of the motor 52 and the motors of the rotors 74A, 74B, 74C, 74D, 85, 89. Mathematically, this looks like a matrix multiplication between the mixer matrix and the controller outputs, as shown in Equation 31:Attorney Docket No.: MIT 25110 PCT | 88212-425635 ^^ ^^ ⋯ ^^Θ^^ ^ଶ ^^^⋮^^ଶ^ ⋱ ⋮ξ^^൩ ൌ ൦ ൪ ^⋮ ൩ , (31)where Θ1, ^ ^ ^ Θnis the set of motor commands, including speed and position, and ξ1, ^ ^ ^ , ξm is the set of controller outputs, such as Fz,thrust, τroll, τpitch, τyaw, and Fx,thrust. The indices of the mixer matrix (a11, ^ ^ ^ , anm) are determined by geometric inspection of the system configuration. Given the physical configuration of aircraft 10 varies depending on the state the aircraft 10 is in, the mixer matrix is unique across almost all states, as will be described in further detail below. In the case of aircraft 10, the motor command mappings are summarized in Table 2 below. FIG.47 and FIG.48 labels the body fixed coordinate frame and motor orientations on the aircraft 10 as related to Table 2, showing the x, y, and z axes. Motor Command Variable Associated Control Surface ω1Quadcopter 1TABLE 2

[0274] A Moore State Machine, also referred to simply as the state machine 128, defined by the output of the state machine 128 depending only on the current state of the system and aircraft 10, can be used for simplicity and clarity in terms of the design. The states in the Moore State Machine 128 are shown in FIG.49. In FIG.49, lines 130 depict safety transitions, lines 132 depict transitions that cross flight modes or prepare the vehicle to crossAttorney Docket No.: MIT 25110 PCT | 88212-425635 a flight mode, and lines 134 depict operational transitions that do not cross flight modes, as will all be described in detail below. Within each state, feedforward linearization strategies and unique mixer matrices (see Table 3) are implemented to integrate with the greater controller architecture outlined above and shown in FIG.46. The following disclosure reviews the standard control scheme used in each state, the key considerations and different states of the state machine 128, as well as any specific linearization strategies used in each state. A person skilled in the art will understand that the various states of the control scheme described below, including their related transition conditions and mixer matrices as well as other aspects thereof, is merely exemplary, and that other states, subsets of states, and other combinations of the same may be utilized based on the control requirements of the aircraft. State Controller Architecture Mixer Matrix Operational States Kill Disarmed - - Armed Quadcopter States ω1 –1 –1 1 1 0 ω2 –1 1 –1 1 0 ω3 –1 1 1 –1 0 Fz, thrust ω4 –1 –1 –1 –10 τroll Quadcopter Hover Multicopter ω5 = 0 0 0 0 0 τpitch ^6 0 0 0 0 0 τyaw ω7 00 0 0 0 Fx, thrust ^8 00 0 0 0 ϰslider 00 0 0 0 VTOL / Hover States ω1 –1 –1 1 1 0 ω2 –1 1 –1 1 0 Rotor Spin Up ω3 –1 1 1 –1 0 Fz, thrust VTOL Multicopter ω4 –1 –1 –1 –1 0 τroll Rotor Accele ω5 = ration 0 0 0 0 0 τpitch ^6 0 0 0 0 0 τyaw ω7 0 0 0 –1 0 Fx, thrust ^8 0 0 0 0 0Attorney Docket No.: MIT 25110 PCT | 88212-425635 State Controller Architecture Mixer Matrix ϰslider 0 0 0 0 0 Forward Flight States ω1 0 0 0 0 0 ω2 0 0 0 0 0ω2 –1 1 –1 1 0 ω3 –1 1 1 –1 0 Fz, thrust Rotor Deceleration ω4 –1 –1 –1 –1 0 τroll Preparation Multicopter ω5 = 0 0 0 0 0 τpitch ^6 0 0 0 0 0 τyaw ω7 0 0 0 0 0 Fx, thrust ^8 0 0 0 0 0 ϰslider 0 0 0 0 0 ω1 –1 –1 1 1 0 ω2 –1 1 –1 1 0 ω3 –1 1 1 –1 0 Fz, thrust ω4 –1 –1 –1 –1 0 τroll Rotor Deceleration Multicopier ω5 = 0 0 0 0 0 τpitch ^6 0 0 0 0 0 τyaw ω7 0 0 0 0 0 Fx, thrust ^8 0 0 0 0 0 ϰslider 0 0 0 0 0 ω1–1 –1 1 1 0 ω2–1 1 –1 1 0 ω3–1 1 1 –1 0 Fz, thrustForward Flight ω4–1 –1 –1 –1 0 τrollPreparation Multicopier ω5 = 0 0 0 0 0 τpitch ^60 0 0 0 0 τyawω7 0 0 0 0 0 Fx, thrust ^8 0 0 0 0 0 ϰslider 0 0 0 0 0 Con –1 –1 1 1 0 ω2 –1 1 –1 1 0 ω3 –1 1 1 –1 0 Fz, thrustForward Acceleration Hybrid ω4 = –1 –1 –1 –1 0 τrollω5 0 0 1 –1 1 τpitch^6 0 –1 0 0 0 τyawω7 0 0 1 1 1 Fx, thrustAttorney Docket No.: MIT 25110 PCT | 88212-425635 State Controller Architecture Mixer Matrix ^80 1 0 0 0 ϰslider 0 0 –1 0 0 Backward Transition States ω1 –1 –1 1 1 0 ω2 –1 1 –1 1 0 ω3 –1 1 1 –1 0 Fz, thrust ω4 –1 –1 –1 –1 0 τroll Forward Deceleration Hybrid ω5 = 0 0 1 –1 1 τpitch ^6 0 –1 0 0 0 τyaw ω7 0 0 11 1 Fx, thrust ^8 0 1 0 0 0 ϰslider 0 0 –1 0 0 ω1 –1 –1 1 1 0 ω2 –1 1 –1 1 0 ω3 –1 1 1 –1 0 Fz, thrust ω4 –1 –1 –1 –1 0 τroll Forward Braking Hybrid ω5 = 0 0 1 –1 0 τpitch ^6 0 –1 0 0 0 τyaw ω7 0 0 1 1 0 Fx, thrust ^8 0 1 0 0 0 ϰslider 0 0 –1 0 0 TABLE 3

[0275] It is noted that the states and transitions shown in FIGS.49-60 may also be referred to herein as follows. The operational states of kill state, disarmed state, and armed state may be referred to as operation safety states. The Quad Hover state, the Rotor Acceleration state, the Rotor Spin Up state, the VTOL state, the Deceleration Prep state, the Rotor Deceleration state, the Forward Flight Prep state, and the Forward Flight state may be referred to as fight states.

[0276] It is further noted that FIGS.49-60 make reference to properties of the states and transitions, in particular those used to instruct the state machine 128 to move between states and transitions. Specifically, the variable, "btn_state" is a variable used to instruct the state machine 128 which mode the user is selecting via a user input, including “0” being an initialization state, “1” being a quadcopter hover, “2” being VTOL flight, and “3” being forward flight. The variable, “arm_state" is a variable used to arm (activate) the various motors and rotors 52, 74A, 74B, 74C, 74D, 85, 89, including “0” being disarm (removes signal from the motors and rotors so that the motors and rotors will not spin even if a command is sent), and “1” being armed (motors will respond to a signal). The variable,Attorney Docket No.: MIT 25110 PCT | 88212-425635 "kill_state" is a variable used to store the safety switch value, including “0” being kill switch disabled, and “1” being kill switch enabled (i.e., kill all motors). The variable, “cb_speed" is a commanded speed of the counterbalance motors 85, 89. The variable, “rotor_accel" is an estimated acceleration of the motor 52. The variable, "x velocity" is a forward speed of the aircraft 10.

[0277] The operational states are the set of states that are necessary for the state machine 128 to operate safely and effective. Namely the states include a kill state, disarmed state, and armed state. Because these states are purely operational and the motors of the aircraft 10 are inactive, there is no control architectures or mixer matrices associated. Each of the operational states are covered in further detail further below. The transition conditions from the operational states are depicted in FIGS.50-52.

[0278] The quadcopter state is the state where the vehicle is operating only as a quadcopter, utilizing the four planar rotors 74A, 74B, 74C, 74D to maneuver and achieve stable flight. Although quadcopter operation is contrary to the goal of creating a stop-rotor aircraft 10, this functionality was included primarily for safety and development of the aircraft 10. In the event that VTOL modes are unstable, the system can transition into quadcopter mode, which is a well-studied and robust configuration. In the quadcopter state, the aircraft 10 is utilizing the pre-canned PX4 multicopter architecture. As such, the main PX4 controller is in standby: the counterbalance rotors 85, 89 are disabled (i.e., ω5= ω7= 0) and held at a constant angle (i.e., θ6 = 0 and θ8 = 180°), and the main rotor is in position hold mode (i.e., θ9= ω9= 0).

[0279] Operating as a quadcopter, the user can control velocity in x and y, position in z, and yaw rate. As with typical quadcopter systems, the velocity in x and y are coupled with pitch and roll respectively: a negative pitch angle induces a positive velocity in x, while a positive roll angle induces a positive velocity in y. As such, roll and pitch cannot be directly controlled by the user. The PX4 controller provides desired Fthrust, τroll, τpitch, and τyaw.

[0280] A transition condition for the quadcopter state, also referred to herein as the Quad Hover state, is described in this paragraph. It is noted that the various “states” described herein refer to those illustrated in the states of FIGS.49-60. The Quad Hover state is reachable from the armed state and the Forward Flight state by selecting quadcopter mode on the mode selection channel. Similarly, from the Quad Hover state, both the RotorAttorney Docket No.: MIT 25110 PCT | 88212-425635 Acceleration state and Forward Flight Preparation state are reachable by selection of VTOL mode and forward flight mode respectively, as depicted in FIG.53. Furthermore, the kill state is reachable by engaging the kill switch.

[0281] A mixer matrix for the quadcopter state is described in this paragraph and shown in Table 3. To generate a positive Fthrust, each quadcopter motor 74A, 74B, 74C, 74D should reduce speed. To generate a positive τroll, motors 2 and 3 must increase speed, while motors 1 and 4 must decrease speed using the motor and coordinate system convention provided in FIG.48. Similarly, to generate a positive τpitch, motors 1 and 3 must increase speed, while motors 2 and 4 must decrease speed. Finally, to generate a positive τyaw, motors 1 and 2 must increase speed, while motors 3 and 4 must decrease speed.

[0282] The VTOL states encompass the set of states that bring the vehicle into and maintain a stable hover with the central lifting surface active including the Rotor Spin Up state, the main VTOL state, and the Rotor Acceleration state, as shown in FIGS.49-60. The VTOL states include VTOL flight of the aircraft 10 as described above with regard to the orientation of the wings 36, 38 as well as the position of the main rotor assembly 30 and the motor assembly 50.

[0283] Similar to the quadcopter states, Fthrust, τroll, τpitch are commanded with the quadcopter platform (motors 74A, 74B, 74C, 74D) in the same way described in in the paragraphs above with regard to motors 1, 2, 3, and 4; however, yaw is commanded differently due to the change in physical configuration of the system. In the Rotor Spin Up state, the counterbalance rotors 85, 89 are positioned such that they can only command a negative yaw torque on the system (i.e., the servos 83, 87 are driven to the extreme position, which is the position facing the aft direction as shown in FIG.13 and FIG.15, to yield a negative yaw torque). Because the main rotor assembly 30, in particular the wings 36, 38 driven in rotation by the motor 52, is accelerating, a positive yaw torque can be achieved by turning off the counterbalance motors 85, 89 and the resulting angular acceleration on the main rotor assembly 30 will generate a positive torque on the main body system. Note that in all cases, θ6 and θ8 are set to 0° and 180°, respectively.

[0284] A mixer matrix for the VTOL states in general will be described in this paragraph and shown in Table 3. In all VTOL modes, the quadcopter architecture is utilized in the same way as described in the quadcopter description above, and therefore the mixer matrixAttorney Docket No.: MIT 25110 PCT | 88212-425635 components associated with the quadcopter motors are the same. Additionally, the speed of the counterbalance rotors 85, 89 is modulated with respect to desired yaw torque. If a negative yaw torque is desired, the counterbalance motor 85A, 89A speeds are increased. Conversely, if a positive yaw torque is desired, the counterbalance motor 85A, 89A speeds are decreased.

[0285] Each individual VTOL state is described in greater detail here. The Rotor Spin Up state accelerates the main rotor assembly 30 to a steady speed from the stopped position while the vehicle is on the ground. This state was included to reduce the influence of the non-linear relationship between main rotor assembly 30 speed and lift while trying to control altitude in VTOL as described in the analytical models above. It is noted that the aircraft 10 remains on the ground for the entirety of the time spent in this state. As a result, there are no control inputs the user can provide to the system: the controller 110 aims to maintain zero x and y velocity, yaw rate, and altitude.

[0286] As the rotor accelerates, the rate of acceleration is limited to 20 rad / s2. In this state, the acceleration of the rotor arotoris computed as the discrete time derivative of rotor speed ωrotor(t), as shown in Equation 32: ^^ ௧ ^^^^ ൌௗ ^^ ^^ௗ௧^^^^௧^^^^^^. (32)The torque due to the^^^^௧^^^^^^ ൌ ^^^^௧^^ ൈ ^^^^௧^^^^^^, (33)where Irotor is estimated to be approximately 0.01 kg ^ m2. The torque from drag on the main rotor assembly 30 is approximated as shown in Equation 34: ^^ ^ ௗ^^^ൌ ଶ ^^^^ௗ^^^^^^^^ଶ^^^^௨^^^^^ଶ^௧^^, (34)where the aerodynamic distance between the axis of rotation and the center of pressure 99A of the wings 36, 38, taken to be 0.1 m. The total yaw torque in the main body system of the aircraft 10 is then computed as shown in Equation 35: ^^௬^௪,௧^௧^^ ൌ ^^௬^௪,^^^௧^^^^^^ െ ^^௬^௪,^^^^௧^^^^௧^^ െ ^^ௗ^^^, (35)Attorney Docket No.: MIT 25110 PCT | 88212-425635 where τyaw,controlleris the computed yaw torque from the closed loop controller and Kyaw,constis 1.15, which was computed computationally using the Simulink framework, as described below. The τyaw,totalvalue is then fed through the mixer 120 and scaled by 1,600 to map into the motor speed domain.

[0287] A transition condition of the Rotor Spin Up state is described in this paragraph. This state is reachable only from the armed state when the channel selection mode is set to VTOL. Besides the kill state, the only other reachable state is the VTOL mode, which the aircraft 10 transitions into automatically once the main rotor assembly 30 reaches the desired speed of 80 rad / s and the rotor acceleration is zero, as summarized in FIG.54.

[0288] The VTOL state is the primary operation state with the motor 52 and main rotor assembly 30 being active. In this state, the aircraft 10 is operating at theoretically maximally efficient hover as the total disk loading of the aircraft 10 is minimized. In the VTOL state, velocity in x and y, position in z, and yaw rate can be controlled by the user. Similar to the quadcopter state, the coupling between pitch and x velocity and roll and y velocity limits pitch and roll from being controlled explicitly by the user.

[0289] It is noted that the mixer matrix is the same as the other VTOL states, as shown in Table 3, and feedforward linearization is performed by compensating for the drag torque. Because the main rotor assembly 30 is held at a constant speed, the compensated yaw torque is computed according to Equation 36: ^^௬^௪,௧^௧^^ ൌ ^^௬^௪,^^^௧^^^^^^ െ ^^ௗ^^^, (36)where τyaw,controller is the desired yaw torque computed from the closed loop controller and τdrag is computed with Equation 34.

[0290] A transition condition of the VTOL state is described in this paragraph. The VTOL state can only be entered from the Rotor Spin Up state and the Rotor Acceleration state by selecting the VTOL mode on the mode selection channel. From this mode, the Rotor Deceleration Preparation state is reachable by selecting any mode that is not VTOL and the kill state is reachable by engaging the kill switch, as summarized in FIG.55.

[0291] In the Rotor Acceleration state, the aircraft 10 starts from a hover with the main rotor assembly 30 at rest and accelerates the main rotor assembly 30 to the final speed of 80Attorney Docket No.: MIT 25110 PCT | 88212-425635 rad / s. It is noted that the Rotor Acceleration state differs from the Rotor Spin Up state due to the starting configuration of the aircraft 10: in the Rotor Spin Up state, the aircraft 10 starts on the ground at rest, while in the Rotor Acceleration state, the aircraft 10 is in a stable hover. Counter torque during main rotor assembly 30 acceleration is provided by the counterbalance rotors 85, 89. Similar to the Rotor Spin Up state, the Rotor Acceleration state compensates for torque induced from main rotor assembly 30 acceleration and main rotor assembly 30 drag by adjusting the desired yaw torque subject to Equation 35.

[0292] A transition condition of the Rotor Acceleration state is described in this paragraph. The Rotor Acceleration state is reachable from the Quad Hover state when VTOL is chosen on the mode selection channel. From the Rotor Acceleration state, the kill state is reachable by engaging the kill switch and the VTOL state is reachable once the main rotor assembly 30 has reached terminal speed and the main rotor assembly 30 acceleration is zero. These reachable states are summarized in FIG.56.

[0293] The VTOL to Forward Flight transition states include all the states that takes the system from the main rotor assembly 30 spinning at steady state in VTOL to cruising in Forward Flight. As shown in the analytical models above, applying some acceleration to the main rotor assembly 30 takes both the yaw and altitude dynamics from 1 / s2systems to 1 / s4. Furthermore, physical reconfiguration should take place to decelerate the main rotor assembly 30 and utilize the counterbalance assembly 80. However, to reconfigure the counterbalance direction without destabilizing the system, the counterbalance assembly 80 should be temporarily disabled. Each state that the system transitions through, as well as the controllers (i.e., controller 110 including the PID and TECS), are described below. It is noted that the aircraft 10 configuration, and therefore mixer matrices, differ between each state, as shown in Table 3.

[0294] To fully decelerate the main rotor assembly 30, the counterbalance assembly 80 must reverse directions to provide torque in the appropriate direction, which requires that the counterbalance rotors 85, 89 be temporarily disabled. When disabled, the yaw dynamics are reduced as shown in Equation 37: ^^^ ^^ௗ௬^^^^^ௗ௬^^^^ ൌ ^^^^௧^^ ∙ ^^^^^௧^^^^^^ ^^^^^ௗ^^^^^^^ଷ^^^^௧^^^^^^ଶ. (37)Attorney Docket No.: MIT 25110 PCT | 88212-425635 If the main rotor assembly 30 speed is held constant, the plant would eventually diverge due to the torque from drag. If the main rotor assembly 30 is decelerated to match the torque acting on the main body system, the plant is considered marginally stable. In other words,for ^^^ (t) = 0, Equation 38 is derived: body^^^ ^^^^ ൌ^ ^^௧^^ଶூ^^^^^^^^^ௗ^^^^^^^ଷ^^^^௧^^^^^^ଶ. (38) However, without aIf, instead, the main rotor assembly 30 acceleration was driven by a combination of drag compensation and controller output such that Equation 39, as shown below, is held true, then the closed loop transfer function is reduced to the form of Equation 14 where the aerodynamic coefficients are summarized in Table 1. Therefore, in this state, the speed of the main rotor assembly 30 is controlled via the numerical solution to the differential Equation 39. భయ మି ఘ^ ^ ^ ఠ ^௧^ ା^^௧^^ ^^^^^^^^మ^ ^ ^^^ ^^ ൌ. (39)^^௧^^ூ^^^^^

[0295] Transition in this paragraph. The Deceleration Preparation state is only reachable from the VTOL state as the aircraft 10 only needs to prepare to decelerate the aircraft 10 from VTOL. The aircraft 10 transitions from VTOL into Deceleration Preparation when either Quadcopter or Fixed Wing modes are selected. As this state rotates the counterbalance rotors 85, 89 into the correct orientation to decelerate the main rotor assembly 30, from this state, only the main rotor assembly 30 deceleration state and kill state are reachable. To reach the Rotor Deceleration state from this state, the counterbalance rotors 85, 89 are to be in the appropriate position and as rest, so the counterbalance rotor 85, 89 speed should be zero. Furthermore, to transition to the kill state, the kill switch must be engaged. The transition conditions to the reachable states are summarized in FIG.57.

[0296] A mixer matrix for the VTOL to Forward Flight transition states will be described in this paragraph and shown in Table 3. While the counterbalance rotors 85, 89 are rotating to the preferred orientation to allow the main rotor assembly 30 to decelerate, the speed and angle should not receive commands from the controllers. As such, only the quadcopter 70Attorney Docket No.: MIT 25110 PCT | 88212-425635 architecture is active. In addition to the closed loop controlled motor commands, the angles of the counterbalance rotors 85, 89, θ6 and θ8, are set to 180° and 0°, respectively.

[0297] Once the counterbalance assembly 80 has fully reversed, the main rotor assembly 30 can then be actively braked or slowed. Similar to the insights provided in description of the analytical models above, for main rotor assembly 30 acceleration, the same conclusions can be drawn for main rotor assembly 30 deceleration: so long as the main rotor assembly 30 is decelerating at a constant value and feedforward linearization is being performed subject to the disturbance provided in Equation 17, the additional controller only needs to stabilize the closed loop transfer function provided in Equation 14.

[0298] As the main rotor assembly 30 decelerates, the deceleration rate is limited to 20 rad / s2. The main rotor assembly 30 deceleration and torque are computed similar to the main rotor assembly 30 acceleration states using Equations 32 and 33, respectively. The total yaw, in addition to the controller requested yaw, is taken to be of similar to form to Equation 35, but it is noted that τrotor will have an opposite sign due to main rotor assembly 30 deceleration rather than acceleration.

[0299] Transition conditions of the Rotor Deceleration state are described in this paragraph. The main rotor assembly 30 deceleration state is only reachable from the Deceleration Preparation state. This is because to properly decelerate the main rotor assembly 30 from VTOL, the counterbalance rotors 85, 89 should be fully rotated, which means that the rotors 85, 89 are rotated to the positions opposite of those shown in FIG.13 and FIG.15 (i.e., the first rotor 85 faces the aft direction and the second rotor 89 faces the forward direction), this also being opposite of the direction of rotation of the main rotor assembly 30. This rotation is managed by the Deceleration Preparation state, as discussed above. The aircraft 10 can transition into the Rotor Deceleration state once the counterbalance rotors 85, 89 have come to rest in the desired orientation. From this state, the Forward Flight Preparation, arm, and kill states are reachable, depending on the value chosen by the user via the mode selection, arm, and kill channels. As summarized in FIG.58, if the kill switch is engaged, the aircraft 10 will transition into the kill state. If any mode but Forward Flight is selected, the aircraft 10 will transition into the arm state once the main rotor assembly 30 has come to rest and automatically transitions into the selected mode (e.g., Quad Hover, Disarmed). If the user has selected the Forward Flight state, once the mainAttorney Docket No.: MIT 25110 PCT | 88212-425635 rotor assembly 30 has come to rest, the aircraft 10 will transition into the Forward Flight Preparation state.

[0300] A mixer matrix for the Rotor Deceleration will be described in this paragraph and shown in Table 3. The counterbalance rotors 85, 89 are used to provide the majority of yaw stabilization and deceleration torque. However, the multicopter 70 architecture is still active and is increasing thrust to accommodate for loss of lift on the main rotor assembly 30. Similar to the Rotor Deceleration condition, θ6and θ8are held at 180° and 0°, respectively.

[0301] Once the main rotor assembly 30 has come to a complete stop and stowed, the orientation of the aircraft 10 should be altered to prepare for forward flight. First, the counterbalance rotors 85, 89 are rotated to orient the motors 85A, 89A forward to produce thrust, as opposed to counter torque in yaw. Second, the linear servo 62 is engaged to push back the main rotor assembly 30 and motor assembly 50 to position the center of pressure 99A of the wings 36, 38 aft the center of gravity 99C of the aircraft 10, as shown in FIG.21. Finally, the wings 36, 38 are configured so the leading edges 36A, 38A are both on the front side of the aircraft 10. In the case where a symmetric wing flips, the servo 41 rotates the second wing 38 approximately 190°. Besides the configuration changes occurring to orient the aircraft 10 for forward flight, the controls of the aircraft 10 in this state are largely the same as the quadcopter controls outlined above with regard to the “quadcopter state.”

[0302] Transition conditions of the Forward Flight Preparation state are described in this paragraph. The Forward Flight Preparation state is reachable from the Quad Hover state, the Rotor Deceleration state, and the Forward Flight state. From the quadcopter state and the Rotor Deceleration state, the aircraft 10 will prepare to transition into forward flight from VTOL. In these cases, the speed of the main rotor assembly 30 is zero and the aircraft 10 should be in a stable hover to transition into the Forward Flight Preparation state. From the Forward Flight state, the aircraft 10 should be moving below 20 m / s to ensure the altitude dynamics are not driven unstable by the multicopter 70 turning on. From this state, the kill state, Forward Flight state, and Quad Hover states are reachable as summarized in FIG.59. The kill state is reachable by engaging the kill switch, the Forward Flight state is reachable when the user has selected forward flight mode, the counterbalance rotors 85, 89 have reached the forward flight orientation (both facing forward), and the aircraft 10 is moving above 20 m / s. Finally, the quadcopter state is reachable when the user has selected any other mode from forward flight.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0303] A mixer matrix of the Forward Flight Preparation state will be described in this paragraph and shown in Table 3. As the aircraft 10 prepares for forward flight, the counterbalance rotors 85, 89 are both rotating to the forward direction and the linear servo 62 is moving to position the wings 36, 38 aft of the center of gravity 99C. As such, none of these motors (i.e., rotor motors 85A, 89A, motor 52) should receive input from the controllers, thereby resulting in the same mixer matrix for pure quadcopter flight as described above.

[0304] The Forward Flight state encompasses the operation of the aircraft 10 where only the main rotor assembly 30 is providing upwards passive lift and the quadcopter 70 architecture is fully disabled. Due to the quadcopter 70 architecture being turned off, the system is only able to transition into the Forward Flight state once a certain velocity has been reached and the wings 36, 38 are generating adequate passive lift.

[0305] In this state, the user has control over the position in z, velocity in x, and velocity in y. As with standard fixed wing aircraft, there is a coupling between the pitch and position in z, as well as the roll and yaw angles and the velocity in y. As such, the roll, pitch, and yaw are not directly controllable by the user. This coupling is managed in the controls using a TECS, as described above.

[0306] Transition conditions of the Forward Flight state are described in this paragraph. The Forward Flight state is only reachable from the Forward Flight Preparation state. Similarly, other than the kill state which is reachable by engaging the kill switch, the only other reachable state is the Forward Flight Preparation state. To enter the Forward Flight Preparation state, the user must specify any mode selection that is not forward flight and the aircraft 10 should have a velocity below 20 m / s so the quadcopter 70 architecture can turn on without destabilizing the altitude of the aircraft 10. The transition conditions out of the Forward Flight state are summarized in FIG.60.

[0307] A mixer matrix of the Forward Flight state will be described in this paragraph and shown in Table 3. The mixer matrix of this state fully removes the control on the quadcopter 70 architecture. Instead, the roll of the aircraft 10 is controlled by modulating the angles of the counterbalance rotors 85, 89, the pitch of the aircraft 10is modulated by the changed center of pressure 99A and the speed of the counterbalance rotors 85, 89, the yaw of the aircraft 10 is modulated by alternating the speed of the counterbalance rotors 85, 89, and theAttorney Docket No.: MIT 25110 PCT | 88212-425635 thrust of the aircraft 10 is modulated by the speed of the counterbalance rotors 85, 89. It is noted that the control surfaces differ from a transitional fixed wing vehicle, which typically contains ailerons, elevators, and rudders.

[0308] The Forward Flight to VTOL transition is handled without explicit states, but by the transition conditions between the Forward Flight state, the Forward Flight Preparation state, the Quad Hover state, and the Rotor Acceleration state. In essence, once the aircraft 10 is below a certain forward velocity, it can engage the quadcopter 70 architecture and come to a stationary hover. Once in a hover, the system can transition into the quadcopter state where the aircraft 10 configuration is positioned for VTOL. Once the aircraft 10 configuration has stabilized, the aircraft 10 can then enter the Rotor Acceleration state. Once the main rotor assembly 30 has reached the terminal velocity, the system automatically transitions into VTOL.

[0309] In some embodiments, the Rotor Deceleration Preparation state, the Rotor Deceleration state, and the Forward Flight Preparation state may be referred to as forward transition states. The Rotor Deceleration Preparation state may be a state in which counterbalance rotors 85, 89 of the aircraft 10 are temporarily disabled and reverse directions, as described above (i.e., facing the opposite directions as those shown in FIG.13 and FIG.15). The Rotor Deceleration state may be a state in which the main rotor assembly 30 is brought to a stop. The Forward Flight Preparation state may be a state in which the wing 38 is rotated to a forward flight position and the main rotor assembly 30 is moved to the aft position so as to position the center of pressure 99A aft of the center of gravity 99C.

[0310] In some embodiments, the Quad Hover state and the Rotor Acceleration states may be referred to as backward transition states. The Quad Hover state may be a state in which the wing 38 is rotated to a VTOL flight position and the main rotor assembly 30 is moved to the forward position such that the center of pressure 99A is approximately aligned with the center of gravity 99C. The counterbalance rotors 85, 89 may be brought back to the VTOL positions shown in FIG.13 and FIG.15 in this state. The Rotor Acceleration state may be a state in which the main rotor assembly 30 is brought to a nominal operating speed for VTOL flight, as described herein.

[0311] The various states, transition between the states, and control of the aircraft 10 in these states encompass the main aspects of the controller 110. The effectiveness of theAttorney Docket No.: MIT 25110 PCT | 88212-425635 controller 110 in controlling and stabilizing the aircraft 10 during flight, and in particular during VTOL to forward flight transitions and vice versa, was validated via deployment of the controller 110 across a simulated environment as well as a real world system. The simulated controller validated the overall control architecture, while the deployed controller accommodates nuances of working with real world systems. The simulated controller, simulation results, and deployed control architecture are reviewed in detail below.

[0312] The pipelines shown in FIG.61 was created in Simulink using a Simulink derived controller and the developed Simscape model described above with regard to the computational models and FIG.45. In FIG.61, r is the desired reference input, e is the error, and y is the output of the system. The developed controller utilized the Moore State Machine 128 discussed above and shown in FIGS.49-60, but within each state is either a multicopter 70 controller for the VTOL and transition states or a TECS controller for the fixed wing states.

[0313] The simulation pipeline shown in FIG.61 was used to verify the stability of the disclosed controls scheme. The successful simulation of aircraft 10 through transition is shown in FIG.62. In the simulation, the aircraft 10 starts at rest (shown as (a) in FIG.62) and spins up the main rotor assembly 30 and takes off into a stable hover (shown as (b) in FIG.62). Once transition is initiated, the aircraft 10 was simulated to prepare for forward flight by holding a stable hover while flipping the second wing 38 to the appropriate direction (shown as (cb) in FIG.62), properly orienting the counterbalance rotors 85, 89 for forward flight (shown as (cc) in FIG.62), and shifting the center of pressure 99A aft of the center of gravity 99C (shown as (cd) in FIG.62). Once the transition is complete, the aircraft 10 was simulated to enter a stable forward flight path (shown as (d) of FIG.62).

[0314] The aircraft 10 response to varying commanded values is depicted in FIGS.63A- 63F for the multicopter 70 operation and FIGS.64A-64C for the fixed wing operation. All multicopter 70 angular controllers (i.e., roll, pitch, and yaw) are able to reach a set-point at a bandwidth of 0.2 Hz or faster rise time, with the yaw controller performing the slowest and the roll controller performing the fastest, as can be seen in FIGS.63A-63C. The x − y position and velocity controllers each exhibit overshoot, perhaps due to the relatively slow roll and pitch controller rates compared to a standard PID tuned controller, as can be seen in FIGS.63D-63F.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0315] In fixed wing flight, the only PID controllers present are the pitch controller, airspeed controller, and altitude controller. The pitch controller operates at approximately 1 Hz (FIG.64A), the airspeed controller at approximately 5 Hz (FIG.64B), and the altitude controller at approximately 0.5 Hz (FIG.64C). In summary, the Simulink-Simscape simulation pipeline provided a proof of concept for the disclosed controls architecture described above. Specifically, it proved that the controllers and feedforward linearization stabilized the system through forward and backward transition (i.e., VTOL flight to forward flight and backwards from forward flight to VTOL flight).

[0316] In deploying the real-world controller to validate the controller 110, the fast and slow dynamics of the system are run on two separate flight controllers. Specifically, the quadcopter 70 system, which handled stabilization of the system, ran on the pre-canned PX4 architecture for a Standard VTOL, and the second flight controller consisted of the Moore State Machine 128 for managing the stop-rotor components and the transition of those components between flight modes. The total deployed control system is summarized in FIG. 65, where Pixhawk 1 is running the pre-canned PX4 Standard VTOL controller and Pixhawk 2 has deployed the custom developed state machine 128 controller to run at 250 Hz. A person skilled in the art will understand that other configurations of control systems may be utilized that implement the control scheme and other aspects thereof as described herein. For example, only a single flight controller as opposed to the two Pixhawk controllers may be utilized, or in other examples, Pixhawk controllers, the PX4, and other components may be replaced with other known similar components that would be applicable to the disclosed embodiments.

[0317] A person skilled in the art will understand that, to operate in the PX4 ecosystem, the deployed Simulink controller should receive accurate state estimate, write motor commands, and read the desired states from a user input. To achieve this, the complete Simulink environment is shown in FIGS.66A-66D. It is noted that the state machine 128 described above is shown in FIG.66C. In addition to the state machine 128, there is an RC Input Processing, shown in FIG.66A, Read and Write uORB Message, shown in FIG.66B, PWM Write, shown in FIG.66D, and Output Processing, shown in FIG.66C.

[0318] Another embodiment of a stop-rotor aircraft 210, 210′ is shown in FIGS.67A-71. The stop-rotor aircraft 210, 210′ is similar to the stop-rotor aircraft 10 described herein. Accordingly, similar reference numbers in the 200, 200′ and 300, 300′ series indicate featuresAttorney Docket No.: MIT 25110 PCT | 88212-425635 that are common between the stop-rotor aircraft 210, 210′ and the stop-rotor aircraft 10. The description of the stop-rotor aircraft 10 is incorporated by reference to apply to the stop-rotor aircraft 210, 210′, except in instances when it conflicts with the specific description and the drawings of the stop-rotor aircraft 210, 210′. It is noted that perspective, top, and other views of the stop-rotor aircraft 210, 210′, although not shown herein, would be substantially similar to those shown in FIGS.1-44B, only differing that, for example, the stop-rotor aircraft 210, 210 also include a second component adjustment assembly 360 in addition to the first component adjustment assembly 260, as described below. A person skilled in the art will be able to reference FIGS.1-44B for any clarification regarding the views shown in FIGS.67A- 71. For example, the lines 28A, 28B-28A, 28B in FIG.12 can provide a reference for the cross-sectional views shown in FIGS.67A, 67B, and 71 (i.e., the views shown in FIGS.67A, 67B, and 71 are through the same cross-sectional line as FIG.28A and FIG.28B, with the difference being that FIGS.67A, 67B, and 71 also show the second component adjustment assembly 360 in addition to the first component adjustment assembly 260 (component adjustment assembly 60 in FIG.28A and FIG.28B)).

[0319] As can be seen in FIGS.67A-71, the stop-rotor aircraft 210, 210′ is formed substantially similarly to the stop-rotor aircraft 10. The stop-rotor aircraft 210, 210′ differs from the stop-rotor aircraft 10 described above in that the stop-rotor aircraft 210, 210′ includes, in one configuration, as shown in FIGS.67A-70, a second component adjustment assembly 360 in addition to the first component adjustment assembly 260 for moving a large component of the aircraft 210 such as the battery pack 296A, and in another configuration, as shown in FIG.71, a second component adjustment assembly 360 for moving a large component of the aircraft 210 such as the battery pack 296A that replaces the first component adjustment assembly 260 (i.e., the aircraft 210′ does not include the first component adjustment assembly 260 such that the main rotor assembly 230 and the motor assembly 250 are fixed to the base frame assembly 220.

[0320] As shown in FIGS.67A-70, the aircraft 210 includes the second component adjustment assembly 360 arranged on an underside of the third support frame plate 223C vertically between the battery pack 296A and the third support frame plate 223C. Illustratively, the second component adjustment 360 may be configured as a linear rail system similar to the first component adjustment assembly 60 described above and the first component adjustment assembly 260 shown in FIGS.67A-70. It is noted that, althoughAttorney Docket No.: MIT 25110 PCT | 88212-425635 referred to here as a second component adjustment assembly, the second component adjustment assembly can also be referred to simply as a component adjustment assembly herein. In some embodiments, the first and second adjustment assemblies may be collectively referred to as at least one component adjustment assembly (i.e., the at least one component adjustment assembly includes the first component adjustment assembly and the second component adjustment assembly).

[0321] Specifically, the second component adjustment assembly 360 includes a linear actuator 362. The second component adjustment assembly 360 can further include a main actuator housing 363 and an actuator rod 364 extending outwardly of the main actuator housing 363 and movable relative to the main actuator housing 363. The main actuator housing 363 is coupled to the third support frame plate 223C, in particular at an aft end thereof. The actuator rod 364 extends outwardly from the main actuator housing 363 and is configured to be actuated by known means, such as internal threads and a lead screw arranged within the main actuator housing 363, to extend the rod 364 in the forward direction away from the main actuator housing 363.

[0322] The second component adjustment assembly 360 further includes a carriage 366 slidably mounted on a linear rail 367, as can be seen in FIGS.67A-70. A relatively large component of the aircraft 210, such as, for example, the battery pack 296A, can be coupled to the carriage 366. The linear rail 367 includes a base rail 367A and a central tongue 367B that both extend along the forward-aft direction. The base rail 67A is coupled to an upper surface of the second support frame plate 23B. In some embodiments, the forward end of the linear rail 367 is located at a position such that the carriage 366 can be moved far enough forward to allow the center of gravity 299C of the aircraft 210 (see FIG.70) to be moved forward enough to be sufficiently spaced apart from the center of pressure 299A, even if the center of pressure 299A was located in the position shown in FIG.16 (i.e., a VTOL flight position).

[0323] As can be seen in FIG.69 the tongue 367B is received in a groove 366A of the carriage 366 so as to locate the carriage 366 on the linear rail 367 and enable slidable movement of the carriage 366 relative to the linear rail 367. In some embodiments, the tongue 367B can have a dovetail shape, and the groove 366A has corresponding recesses to receive the dovetail shaped tongue 367B, similar to the linear rail 67 described above. The carriage 366 may further include a lower surface configured to slide along an upper surface of the base rail 367A.Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0324] In operation, the linear actuator 362 actuates the actuator rod 364 so as to slidably move the carriage 366 along the linear rail 367. In this way, because the battery pack 296A is fixedly coupled to the carriage 366, the battery pack 296A can be moved forward and aft relative to the base frame assembly 220, the multicopter assembly 270, the counterbalance rotor assembly 280, and the tail assembly 290. Illustratively, the carriage 366 along with the battery pack 296A can be arranged in an aft position, as shown in FIG.67A, and a forward position, as shown in FIG.67B. It is noted that the aft position may be the same position of the center of gravity 99C shown in FIG.17.

[0325] Because the battery pack 296A is a large, relatively heavy component of the aircraft 210, moving the battery pack 296A forward and aft will affect the center of gravity 299C of the aircraft 210. It is noted that the battery pack 296A may also be referred to herein as a “component” that is movable by the second component adjustment assembly 360. A person skilled in the art will understand that other components may be arranged on the second component adjustment assembly 360, in particular components whose movement forward and aft would sufficiently affect the center of gravity 299C. As non-limiting examples, other components that could be arranged on the carriage 366 could include other electronics such as radios and communication equipment, flight computers, power distribution systems, and other similar components. Such components may also be referred o as “components” that are movable by the second component adjustment assembly 360.

[0326] As can be seen in FIG.67B and FIG.70, the center of gravity 299C of the aircraft 210 can be adjusted by moving the carriage 366 and thus the battery pack 296A forward and aft along the linear rail 367. As a result, the distance between the center of gravity 299C and the center of pressure 299A can be actively adjusted during flight, which provides a great deal of flexibility for stability adjustments. For example, as shown in FIG.67B, the main rotor assembly 230 and the motor assembly 250 can remain in the starting, forward position, while the battery pack 296A is moved to a forward position along the rail 367, thus adjusting the center of gravity 299C forward of the center of pressure 299A. As a result, because the center of pressure 299A is located aft of the center of gravity 299C, stability is improved during forward flight. As another example, as shown in FIG.70, the main rotor assembly 230 and the motor assembly 250 are arranged in an aft position (which may not be entirely aft along the rail 267, as shown in FIG.70), while the battery pack 296A is moved to a forward position along the rail 367, thus adjusting the center of gravity 299C forward of theAttorney Docket No.: MIT 25110 PCT | 88212-425635 center of pressure 299A to a distance further away from the center of pressure 299A than as shown in FIG.67B. This greater distance may be advantageous in certain flight scenarios. The carriages 266, 366 are movable and lockable at any position along the rails 267, 367, which thus allows for any distance between the center of gravity 299C and the center of pressure 299A ranging from (i) the positions shown in FIG.17 and (ii) the forwardmost position of the carriage 366 (not shown, but a position at which the forward end of the carriage 366 is located at the forward end of the rail 367) and the aftmost position of the carriage 266 (see FIG.21).

[0327] Is noted that, in addition to the first, second, and third flight configurations described above with respect to FIGS.16, 17, and 21, the positions of the batter pack 296A and thus the center of gravity 299C relative to the center of pressure 299A can also be considered flight configurations of the aircraft 10, 210, 210′ in addition to or alternatively to the first, second, and third configurations. For example, the center of gravity 299C being moved forward via movement of the battery pack 296A or other component forward to a forward position along with the main rotor assembly 230 and the motor assembly 250, and thus the center of pressure 299A, being in the VTOL, forward position may be considered an additional flight configuration, or a fourth flight configuration. This can also apply to the configuration of the aircraft 210′ described below in which the main rotor assembly 230 and the motor assembly 250 are fixed relative to the base frame assembly 220. As a further example, the center of gravity 299C being moved forward via movement of the battery pack 296A or other component forward to a forward position along with the main rotor assembly 230 and the motor assembly 250, and thus the center of pressure 299A, being moved rearward into the aft position may be considered an additional flight configuration, or a fifth flight configuration. A person skilled in the art will understand that any number of flight configurations are possible via selective movement of the center of pressure 299A and the center of gravity 299C relative to each other, as is possible with the main rotor assembly 230 and the motor assembly 250, as well as the battery pack 296A or other component, both being movable.

[0328] A person skilled in the art will understand that, similarly to the first component adjustment assembly 60, 260, the second component adjustment 360 may include other means of adjusting the forward-aft location of the carriage 366 and thus the battery pack 296A, such as, for example, pneumatic, hydraulic, and similar actuators. Moreover, otherAttorney Docket No.: MIT 25110 PCT | 88212-425635 means of moving the carriage 366 in the forward-aft direction would be understood by a person skilled in the art, such as, for example, rollers, belt drives, bearings, and the like.

[0329] An alternative configuration of an aircraft 210′ is shown in FIG.71. Unlike the configuration shown in FIGS.67A-70, the aircraft 210′ does not include a component adjustment assembly for the main rotor assembly 230′ and the motor assembly 250′, but instead only includes a component adjustment assembly 360′ for the battery pack 296A′. As such, only the center of gravity 299C is movable forward and aft. Although forward-aft movement of the main rotor assembly 230′ and the motor assembly 250′ is not available, the center of pressure 299A′ can still be made to be aft of the center of gravity 299C′ by moving the battery pack 296A′ forward, while potentially saving on weight, complexity of design, and other factors via the removal of the component adjustment assembly for the main rotor assembly 230′ and the motor assembly 250′.

[0330] FIG.72 is a schematic diagram that shows a non-limiting example of a computing system 400 that can be used to implement the techniques described herein. The computing system 400 includes one or more computing devices (e.g., controller 410), which can be the equivalent of any of the controllers described herein, including the controller 110. The controller 410 is configured to execute any and all steps and actions described above with regard to aircraft 10, 210, 210′ control and stabilization, such as, for example, those shown in FIGS.44-66D.

[0331] The controller 410 can be in wired and / or wireless communication with various peripheral device(s) 480, data source(s) 490, and / or other computing devices (e.g., over network(s) 470). The controller 410 can represent various forms of stationary computers 412 (e.g., workstations, kiosks, servers, mainframes, edge computing devices, quantum computers, etc.) and mobile computers 414 (e.g., laptops, tablets, mobile phones, personal digital assistants, wearable devices, etc.). In some implementations, the controller 410 can be included in (and / or in communication with) various other sorts of devices, such as data collection devices (e.g., devices that are configured to collect data from a physical environment), robotic devices (e.g., devices that are configured to physically interact with objects in a physical environment, such as manufacturing devices, maintenance devices, object handling devices, etc.), vehicles (e.g., devices that are configured to move throughout a physical environment, such as automated guided vehicles, manually operated vehicles, etc.), or other such devices. Each of the devices (e.g., stationary computers, mobileAttorney Docket No.: MIT 25110 PCT | 88212-425635 computers, and / or other devices) can include components of the controller 410, and an entire system can be made up of multiple devices communicating with each other. For example, the controller 410 can be part of a computing system that includes a network of computing devices, such as a cloud-based computing system, a computing system in an internal network, or a computing system in another sort of shared network. Processors of the computing device controller 410 and other computing devices of a computing system can be optimized for different types of operations, secure computing tasks, etc. The components shown herein, and their functions, are meant to be examples, and are not meant to limit implementations of the technology described and / or claimed in this document.

[0332] The controller 410 can include processor(s) 420, memory device(s) 430, storage device(s) 440, and interface(s) 450. Each of the processor(s) 420, the memory device(s) 430, the storage device(s) 440, and the interface(s) 450 can be interconnected using a system bus 460. The processor(s) 420 are capable of processing instructions for execution within the controller 410 and can include one or more single-threaded and / or multi-threaded processors. The processor(s) 420 are capable of processing instructions stored in the memory device(s) 430 and / or on the storage device(s) 440. The memory device(s) 430 can store data within the controller 410, and can include one or more computer-readable media, volatile memory units, and / or non-volatile memory units. The storage device(s) 440 can provide mass storage for the controller 410, can include various computer-readable media (e.g., a floppy disk device, a hard disk device, a tape device, an optical disk device, a flash memory or other similar solid state memory device, or an array of devices, including devices in a storage area network or other configurations), and can provide date security / encryption capabilities.

[0333] The interface(s) 450can include various communications interfaces (e.g., USB, Near-Field Communication (NFC), Bluetooth, WiFi, Ethernet, wireless Ethernet, etc.) that can be coupled to the network(s) 470, peripheral device(s) 480, and / or data source(s) 490 (e.g., through a communications port, a network adapter, etc.). Communication can be provided under various modes or protocols for wired and / or wireless communication. Such communication can occur, for example, through a transceiver using a radio-frequency. As another example, communication can occur using light (e.g., laser, infrared, etc.) to transmit data. As another example, short-range communication can occur, such as using Bluetooth, WiFi, or other such transceiver. In addition, a GPS (Global Positioning System) receiver module can provide location-related wireless data, which can be used as appropriate byAttorney Docket No.: MIT 25110 PCT | 88212-425635 device applications. The interface(s) 450 can include a control interface that receives commands from an input device (e.g., operated by a user) and converts the commands for submission to the processors 420. The interface(s) 450 can include a display interface that includes circuitry for driving a display to present visual information to a user. The interface(s) 450 can include an audio codec which can receive sound signals (e.g., spoken information from a user) and convert it to usable digital data. The audio codec can likewise generate audible sound, such as through an audio speaker. Such sound can include real-time voice communications, recorded sound (e.g., voice messages, music files, etc.), and / or sound generated by device applications.

[0334] The network(s) 470 can include one or more wired and / or wireless communications networks, including various public and / or private networks. Examples of communication networks include a LAN (local area network), a WAN (wide area network), and / or the Internet. The communication networks can include a group of nodes (e.g., computing devices) that are configured to exchange data (e.g., analog messages, digital messages, etc.), through telecommunications links. The telecommunications links can use various techniques (e.g., circuit switching, message switching, packet switching, etc.) to send the data and other signals from an originating node to a destination node. In some implementations, the controller 410 can communicate with the peripheral device(s) 480, the data source(s) 490, and / or other computing devices over the network(s) 470. In some implementations, the controller 410 can directly communicate with the peripheral device(s) 480, the data source(s), and / or other computing devices.

[0335] The peripheral device(s) 480 can provide input / output operations for the controller 410. Input devices (e.g., keyboards, pointing devices, touchscreens, microphones, cameras, scanners, sensors, etc.) can provide input to the controller 410 (e.g., user input and / or other input from a physical environment). Output devices (e.g., display units such as display screens or projection devices for displaying graphical user interfaces (GUIs)), audio speakers for generating sound, tactile feedback devices, printers, motors, hardware control devices, etc.) can provide output from the controller 410. Other kinds of devices can be used to provide for interactions between users and devices. For example, input from a user can be received in any form, including visual, auditory, or tactile input, and feedback provided to the user can be any form of sensory feedback (e.g., visual feedback, auditory feedback, or tactile feedback).Attorney Docket No.: MIT 25110 PCT | 88212-425635

[0336] The data source(s) 490 can provide data for use by the controller 410, and / or can maintain data that has been generated by the controller 410 and / or other devices (e.g., data collected from sensor devices, data aggregated from various different data repositories, etc.). Data can be provided to the controller 410, and the controller 410 can relay and store data regarding the same to and from the various storage and processing devices of the system 400. In some implementations, one or more data sources can be hosted by the controller 410 (e.g., using the storage device(s) 440). In some implementations, one or more data sources can be hosted by a different computing device. Data can be provided by the data source(s) 490 in response to a request for data from the controller 410 and / or can be provided without such a request. For example, a pull technology can be used in which the provision of data is driven by device requests, and / or a push technology can be used in which the provision of data occurs as the data becomes available (e.g., real-time data streaming and / or notifications). Various sorts of data sources can be used to implement the techniques described herein, alone or in combination.

[0337] In some implementations, a data source can include one or more data store(s) 490a. The database(s) can be provided by a single computing device or network (e.g., on a file system of a server device) or provided by multiple distributed computing devices or networks (e.g., hosted by a computer cluster, hosted in cloud storage, etc.). In some implementations, a database management system (DBMS) can be included to provide access to data contained in the database(s) (e.g., through the use of a query language and / or application programming interfaces (APIs)). The database(s), for example, can include relational databases, object databases, structured document databases, unstructured document databases, graph databases, and other appropriate types of databases.

[0338] Examples of the above-described embodiments can include the following: 1. An aircraft, comprising: a base frame assembly; a main rotor assembly including at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, the at least one wing defining a center of pressure; and a component adjustment assembly arranged on the base frame assembly and configured to move at least one component of the aircraft relative to the base frame assemblyAttorney Docket No.: MIT 25110 PCT | 88212-425635 to one or more of move the center of pressure of the at least one wing relative to a center of gravity of the aircraft or move the center of gravity relative to the center of pressure. 2. The aircraft of example 1, wherein the at least one component of the aircraft that is movable by the component adjustment assembly includes the main rotor assembly. 3. The aircraft of example 2, wherein the component adjustment assembly is configured to move the main rotor assembly in a forward-aft direction to move the center of pressure of the at least one wing in the forward-aft direction. 4. The aircraft of any of examples 1 to 3, wherein the component adjustment assembly includes a carriage configured to move in the forward-aft direction relative to the base frame assembly, and wherein the main rotor assembly is coupled to the carriage of the component adjustment assembly such that movement of the carriage relative to the base frame assembly moves the center of pressure of the at least one wing forward and aft relative to the center of gravity of the aircraft. 5. The aircraft of example 4, wherein the carriage is movable between a forward position for VTOL flight of the aircraft in which the center of pressure of the at least one wing is approximately aligned with the center of gravity of the aircraft in the forward-aft direction and an aft position for forward flight of the aircraft in which the center of pressure of the at least one wing is aft of the center of gravity of the aircraft in the forward-aft direction. 6. The aircraft of example 4 or 5, wherein the component adjustment assembly further includes a linear rail arranged on the base frame assembly that extends in the forward-aft direction, and wherein the carriage is slidably movable on the linear rail. 7. The aircraft of example 6, wherein the component adjustment assembly further includes a linear actuator and an actuator rod coupled to the carriage, and wherein the linear actuator is configured to move the actuator rod toward and away from the linear actuator to move the carriage toward and away from the linear actuator along the linear rail.Attorney Docket No.: MIT 25110 PCT | 88212-425635 8. The aircraft of any of examples 4 to 7, further comprising: a motor assembly including a motor coupled to a central housing of the main rotor assembly from which the at least one wing extends, the motor configured to rotate the main rotor assembly, wherein the motor assembly is arranged on the carriage of the component adjustment assembly such that the motor assembly and the main rotor assembly are configured to be moved relative to the base frame assembly via movement of the carriage relative to the base frame assembly. 9. The aircraft of example 8, wherein the at least one wing is rotatable relative to the central housing between a VTOL position in which a leading edge of the at least one wing faces a direction of rotation of the main rotor assembly about the axis and a forward flight position in which the leading edge faces a forward direction of the aircraft. 10. The aircraft of example 9, wherein the at least one wing includes a first wing that is fixed relative to the central housing and a second wing that is rotatable relative to the central housing, and wherein the motor and first and second wings are configured to be rotationally locked in forward flight. 11. The aircraft of example 10, wherein the leading edges of the first and second wings face opposite directions in the VTOL position of the second wing, and wherein the leading edges of the first and second wings face the forward direction in the forward flight position of the second wing. 12. The aircraft of any of examples 9 to 11, wherein the main rotor assembly further includes a transition motor assembly configured to rotate the at least one wing between and including the VTOL and forward flight positions, and wherein the transition motor assembly includes a locking bumper arranged in a fixed position relative to the at least one wing. 13. The aircraft of example 12, wherein the transition motor assembly further includes a locking pin configured to be engaged by the locking bumper during rotation of the at leastAttorney Docket No.: MIT 25110 PCT | 88212-425635 one wing from the VTOL position to the forward flight position, and configured to be disengaged from the locking bumper during rotation of the at least one wing from the forward flight position to the VTOL position. 14. The aircraft of example 13, wherein engagement of the locking bumper with the locking pin during rotation of the at least one wing from the VTOL position to the forward flight position causes the locking pin to enter a hole formed in a rotor plate fixedly disposed on the base frame assembly to rotationally lock the main rotor assembly relative to the base frame assembly. 15. The aircraft of example 13 or 14, wherein disengagement of the locking bumper from the locking pin during rotation of the at least one wing from the forward flight position to the VTOL position causes the locking pin to be removed from a hole formed in a rotor plate fixedly disposed on the base frame assembly to rotationally unlock the main rotor assembly relative to the base frame assembly. 16. The aircraft of any preceding example, further comprising: a counterbalance assembly including a first counterbalance subassembly having a first counterbalance rotor and extending away from a first side of the base frame assembly and a second counterbalance subassembly having a second counterbalance rotor and extending away from a second side of the base frame assembly opposite the first side. 17. The aircraft of example 16, wherein the first and second counterbalance rotors are rotatable between a forward-facing position and an aft-facing position. 18. The aircraft of example 17, wherein, in a VTOL configuration, the first counterbalance rotor is in the forward-facing position and the second counterbalance rotor is in the aft-facing position such that first and second counterbalance rotors face a direction of rotation of the main rotor assembly about the axis such that the first and second counterbalance rotors stabilize a yaw of the aircraft. 19. The aircraft of example 17 or 18, wherein, in a forward flight configuration, the first counterbalance rotor is in the forward-facing position and the second counterbalance rotor is in the forward-facing position such that first and second counterbalance rotors face a forward direction of the aircraft to generate forward thrust to the aircraft.Attorney Docket No.: MIT 25110 PCT | 88212-425635 20. The aircraft of any preceding example, further comprising: a multicopter assembly arranged on the base frame assembly and including a multicopter plate and a plurality of multicopter rotors arranged on the multicopter plate, the plurality of multicopter rotors configured to generate thrust in a downward direction so as to provide lift to the aircraft. 21. The aircraft of example 20, wherein the plurality of multicopter rotors are configured to be rotating in a VTOL mode of the aircraft and in a transition mode of the aircraft between VTOL and forward flight to provide lift to the aircraft, and configured to be shut off in a forward flight mode of the aircraft. 22. The aircraft of example 20 or 21, wherein the plurality of multicopter rotors includes four multicopter rotors. 23. The aircraft of any of examples 20 to 22, wherein the multicopter plate is arranged on the base frame assembly above the center of gravity of the aircraft. 24. The aircraft of any of examples 20 to 23, wherein the multicopter plate is arranged on the base frame assembly above the center of gravity of the aircraft such that a center of thrust of the plurality of multicopter rotors is arranged above the center of gravity. 25. The aircraft of any preceding example, wherein the at least one component of the aircraft that is movable by the component adjustment assembly includes a first component of the aircraft, and wherein the component adjustment assembly is configured to move the first component in a forward-aft direction to move the center of gravity of the aircraft in the forward-aft direction. 26. The aircraft of example 25, wherein the first component is an electronics component of the aircraft. 27. The aircraft of example 25 or 26, wherein the component adjustment assembly includes a carriage configured to move in the forward-aft direction relative to the base frame assembly, and wherein the first component is coupled to the carriage of the component adjustmentAttorney Docket No.: MIT 25110 PCT | 88212-425635 assembly such that movement of the carriage relative to the base frame assembly moves the center of gravity forward and aft relative to the center of pressure of the at least one wing. 28. The aircraft of example 27, wherein the carriage is movable between an aft position for VTOL flight of the aircraft in which the center of pressure of the at least one wing is approximately aligned with the center of gravity of the aircraft in the forward-aft direction and a forward position for forward flight of the aircraft in which the center of pressure of the at least one wing is aft of the center of gravity of the aircraft in the forward-aft direction. 29. The aircraft of example 27 or 28, wherein the component adjustment assembly further includes a linear rail arranged on an underside of the base frame assembly that extends in the forward-aft direction, and wherein the carriage is slidably movable on the linear rail. 30. The aircraft of example 29, wherein the component adjustment assembly further includes a linear actuator and an actuator rod coupled to the carriage, and wherein the linear actuator is configured to move the actuator rod toward and away from the linear actuator to move the carriage toward and away from the linear actuator along the linear rail. 31. A method, comprising: providing a base frame assembly of an aircraft; coupling a main rotor assembly of the aircraft to a component adjustment assembly of the aircraft, the main rotor assembly including at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, the at least one wing defining a center of pressure; arranging the component adjustment assembly on the base frame assembly, the component adjustment assembly being configured to move at least one component of the aircraft relative to the base frame assembly to one or more of move the center of pressure of the at least one wing relative to a center of gravity of the aircraft or move the center of gravity relative to the center of pressure. 32. The method of example 31, wherein the at least one component of the aircraft that is movable by the component adjustment assembly includes the main rotor assembly.Attorney Docket No.: MIT 25110 PCT | 88212-425635 33. The method of example 32, wherein the component adjustment assembly is configured to move the main rotor assembly in a forward-aft direction to move the center of pressure of the at least one wing in the forward-aft direction. 34. The method of example 32 or 33, wherein the component adjustment assembly includes a carriage configured to move in the forward-aft direction relative to the base frame assembly, the method further comprising: coupling the main rotor assembly to the carriage of the component adjustment assembly such that movement of the carriage relative to the base frame assembly moves the center of pressure of the at least one wing forward and aft relative to the center of gravity of the aircraft. 35. The method of example 34, wherein the carriage is movable between a forward position for VTOL flight of the aircraft in which the center of pressure of the at least one wing is approximately aligned with the center of gravity of the aircraft in the forward-aft direction and an aft position for forward flight of the aircraft in which the center of pressure of the at least one wing is aft of the center of gravity of the aircraft in the forward-aft direction. 36. The method of example 35, further comprising: arranging a linear rail of the component adjustment assembly on the base frame assembly, the linear rail extending in the forward-aft direction, wherein the carriage is slidably movable on the linear rail. 37. The method of example 36, further comprising: arranging a linear actuator of the component adjustment assembly on the base frame, the component adjustment assembly further including an actuator rod that is actuatable by the linear actuator; and coupling the actuator rod to the carriage, wherein the linear actuator is configured to move the actuator rod toward and away from the linear actuator to move the carriage toward and away from the linear actuator along the linear rail.Attorney Docket No.: MIT 25110 PCT | 88212-425635 38. The method of any of examples 31 to 37, further comprising: coupling a motor of a motor assembly of the aircraft to a central housing of the main rotor assembly from which the at least one wing extends, the motor being configured to rotate the main rotor assembly; and arranging the motor assembly on the carriage of the component adjustment assembly such that the motor assembly and the main rotor assembly are configured to be moved relative to the base frame assembly via movement of the carriage relative to the base frame assembly. 39. The method of any of examples 31 to 38, wherein the at least one wing is rotatable relative to the central housing between a VTOL position in which a leading edge of the at least one wing faces a direction of rotation of the main rotor assembly about the axis and a forward flight position in which the leading edge faces a forward direction of the aircraft. 40. The method of example 39, wherein the at least one wing includes a first wing that is fixed relative to the central housing and a second wing that is rotatable relative to the central housing, and wherein the motor and first and second wings are configured to be rotationally locked in forward flight. 41. The method of example 40, wherein the leading edges of the first and second wings face opposite directions in the VTOL position of the second wing, and wherein the leading edges of the first and second wings face the forward direction in the forward flight position of the second wing. 42. The method of any of examples 39 to 41, wherein the main rotor assembly further includes a transition motor assembly configured to rotate the at least one wing between and including the VTOL and forward flight positions, and wherein the transition motor assembly includes a locking bumper arranged in a fixed position relative to the at least one wing. 43. The method of example 42, wherein the transition motor assembly further includes a locking pin configured to be engaged by the locking bumper during rotation of the at leastAttorney Docket No.: MIT 25110 PCT | 88212-425635 one wing from the VTOL position to the forward flight position, and configured to be disengaged from the locking bumper during rotation of the at least one wing from the forward flight position to the VTOL position. 44. The method of example 43, wherein engagement of the locking bumper with the locking pin during rotation of the at least one wing from the VTOL position to the forward flight position causes the locking pin to enter a hole formed in a rotor plate fixedly disposed on the base frame assembly to rotationally lock the main rotor assembly relative to the base frame assembly. 45. The method of example 43 or 44, wherein disengagement of the locking bumper from the locking pin during rotation of the at least one wing from the forward flight position to the VTOL position causes the locking pin to be removed from a hole formed in a rotor plate fixedly disposed on the base frame assembly to rotationally unlock the main rotor assembly relative to the base frame assembly. 46. The method of any of examples 31 to 45, further comprising: arranging a first counterbalance subassembly of a counterbalance assembly of the aircraft on a first side of the base frame assembly such that the first counterbalance subassembly extends away from the first side of the base frame assembly, the first counterbalance subassembly including a first counterbalance rotor; and arranging a second counterbalance subassembly of the counterbalance assembly on a second side of the base frame assembly opposite the first side such that the second counterbalance subassembly extends away from the second side of the base frame assembly, the second counterbalance subassembly including a second counterbalance rotor. 47. The method of example 46, wherein the first and second counterbalance rotors are rotatable between a forward-facing position and an aft-facing position. 48. The method of example 47, wherein, in a VTOL configuration, the first counterbalance rotor is in the forward-facing position and the second counterbalance rotor is in the aft-facing position such that first and second counterbalance rotors face a direction of rotation of the main rotor assembly about the axis such that the first and second counterbalance rotors stabilize a yaw of the aircraft.Attorney Docket No.: MIT 25110 PCT | 88212-425635 49. The method of example 46 or 48, wherein, in a forward flight configuration, the first counterbalance rotor is in the forward-facing position and the second counterbalance rotor is in the forward-facing position such that first and second counterbalance rotors face a forward direction of the aircraft to generate forward thrust to the aircraft. 50. The method of any of examples 31 to 49, further comprising: arranging a multicopter assembly of the aircraft on the base frame assembly, the multicopter assembly including a multicopter plate and a plurality of multicopter rotors arranged on the multicopter plate, the plurality of multicopter rotors configured to generate thrust in a downward direction so as to provide lift to the aircraft. 51. The method of example 50, wherein the plurality of multicopter rotors are configured to be rotating in a VTOL mode of the aircraft and in a transition mode of the aircraft between VTOL and forward flight to provide lift to the aircraft, and configured to be shut off in a forward flight mode of the aircraft. 52. The method of example 50 or 51, wherein the plurality of multicopter rotors includes four multicopter rotors. 53. The method of any of examples 50 to 52, wherein the multicopter plate is arranged on the base frame assembly above the center of gravity of the aircraft. 54. The method of any of examples 50 to 52, wherein the multicopter plate is arranged on the base frame assembly above the center of gravity of the aircraft such that a center of thrust of the plurality of multicopter rotors is arranged above the center of gravity. 55. An aircraft, comprising: a base frame assembly; a multicopter assembly arranged on the base frame assembly, the multicopter assembly including a plurality of multicopter rotors configured to generate thrust in a downward direction so as to provide lift to the aircraft; and a main rotor assembly arranged above the multicopter assembly and including at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, wherein the multicopter plate is arranged on the base frame assembly above the center of gravity of the aircraft.Attorney Docket No.: MIT 25110 PCT | 88212-425635 56. The aircraft of example 55, wherein the multicopter assembly includes a multicopter plate, the plurality of multicopter rotors being arranged on the multicopter plate. 57. The aircraft of example 56, wherein the multicopter plate is arranged on the base frame assembly above the center of gravity of the aircraft such that a center of thrust of the plurality of multicopter rotors is arranged above the center of gravity. 58. The aircraft of example 56 or 57, wherein the plurality of multicopter rotors are configured to be rotating in a VTOL mode of the aircraft and in a transition mode of the aircraft between VTOL and forward flight to provide lift to the aircraft, and configured to be shut off in a forward flight mode of the aircraft. 59. The aircraft of any of examples 55 to 58, further comprising: a component adjustment assembly arranged on the base frame assembly and including a carriage configured to move relative to the base frame assembly, wherein the main rotor assembly is coupled to the carriage, and wherein the carriage is configured to be moved relative to the base frame assembly to move the main rotor assembly relative to the base frame assembly and thus move a center of pressure of the at least one wing relative to the center of gravity of the aircraft. 60. A method, comprising: providing a base frame assembly of an aircraft; arranging a multicopter assembly of the aircraft on the base frame assembly, the multicopter assembly including a plurality of multicopter rotors configured to generate thrust in a downward direction so as to provide lift to the aircraft; and arranging a main rotor assembly of the aircraft above the multicopter assembly, the main rotor assembly including at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, wherein the multicopter plate is arranged on the base frame assembly below the center of gravity of the aircraft. 61. The method of example 60, wherein the multicopter assembly includes a multicopter plate, the plurality of multicopter rotors being arranged on the multicopter plate. 62. The method of example 60 or 61, wherein the plurality of multicopter rotors are configured to be rotating in a VTOL mode of the aircraft and in a transition mode of theAttorney Docket No.: MIT 25110 PCT | 88212-425635 aircraft between VTOL and forward flight to provide lift to the aircraft, and configured to be shut off in a forward flight mode of the aircraft. 63. The method of any of examples 60 to 62, further comprising: arranging a component adjustment assembly of the aircraft on the base frame assembly, the component adjustment assembly including a carriage configured to move relative to the base frame assembly; and coupling the main rotor assembly to the carriage, wherein the carriage is configured to be moved relative to the base frame assembly to move the main rotor assembly relative to the base frame assembly and thus move a center of pressure of the at least one wing relative to the center of gravity of the aircraft. 64. A method of operating an aircraft, comprising: operating an aircraft in a VTOL configuration in which at least one component of the aircraft is arranged relative to the base frame assembly such that a center of pressure of at least one wing of a main rotor assembly of the aircraft is approximately aligned with a center of gravity of the aircraft in a forward-aft direction; moving the at least one component relative to the base frame assembly to a forward flight configuration such that the center of pressure of the at least one wing is aft of the center of gravity of the aircraft; and operating the aircraft in a forward flight configuration. 65. The method of example 64, wherein the aircraft includes a base frame assembly, the main rotor assembly including the at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, the at least one wing defining the center of pressure, and at least one component adjustment assembly arranged on the base frame assembly and configured to move at least one component of the aircraft relative to the base frame assembly. 66. The method of example 65, wherein, in the VTOL configuration, a leading edge of the at least one wing faces a direction of rotation of the main rotor assembly about the axis. 67. The method of example 65 or 66, further comprising: transitioning from the VTOL configuration of the aircraft to the forward flight configuration which includes rotating the at least one wing relative to an axis of rotation ofAttorney Docket No.: MIT 25110 PCT | 88212-425635 the at least one wing such that the leading edge faces a forward direction of the aircraft for forward flight. 68. The method of any of examples 64 to 67, further comprising: during the operating of the aircraft in the VTOL configuration, operating first and second counterbalance rotors of the aircraft that each face in the direction of rotation of the main rotor assembly. 69. The method of example 67 or 68, further comprising: during the transitioning from the VTOL configuration of the aircraft to the forward flight configuration, rotating one of the first and second counterbalance rotors to face a forward direction such that both of the first and second counterbalance rotors face the forward direction such that the first and second counterbalance rotors generate forward thrust for forward flight. 70. The method of any of examples 64 to 69, wherein the at least one component of the aircraft that is movable by the at least one component adjustment assembly includes the main rotor assembly, the method further comprising: moving the main rotor assembly via the at least one component adjustment assembly in a forward-aft direction to move the center of pressure of the at least one wing in the forward-aft direction. 71. The method of example 70, further comprising: moving a carriage of the at least one component adjustment assembly to which the main rotor assembly is coupled in the forward-aft direction relative to the base frame assembly such that movement of the carriage relative to the base frame assembly moves the center of pressure of the at least one wing forward and aft relative to the center of gravity of the aircraft. 72. The method of example 71, further comprising: sliding the carriage along a linear rail of the at least one component adjustment assembly arranged on the base frame assembly. 73. The method of any of examples 64 to 72, wherein the at least one component of the aircraft that is movable by the at least one component adjustment assembly includes a first component of the aircraft, the method further comprising:Attorney Docket No.: MIT 25110 PCT | 88212-425635 moving the first component via the at least one component adjustment assembly in a forward-aft direction to move the center of gravity of aircraft in the forward-aft direction. 74. The method of example 73, wherein the first component is an electronics component of the aircraft. 75. The method of example 73 or 74, further comprising: moving a carriage of the at least one component adjustment assembly to which the first component is coupled in the forward-aft direction relative to the base frame assembly such that movement of the carriage relative to the base frame assembly moves the center of gravity of the aircraft forward and aft relative to the center of pressure of the at least one wing. 76. The method of any of examples 64 to 75, utilizing a state machine control scheme to operate the aircraft, the state machine control scheme including: at least one flight state including at least one of a VTOL state, a forward flight state, at least one forward transition state for transitioning from VTOL to forward flight, or a at least one backward transition state for transitioning from forward flight to VTOL; and at least one operational safety state including at least one of a kill state, a disarm state, or an arm state, wherein the control scheme is configured to transition between the at least one flight state and the at least one operational safety state so as to transition the aircraft between VTOL flight and forward flight. 77. The method of example 76, wherein the at least one forward transition state includes at least one of a rotor deceleration preparation state in which counterbalance rotors of the aircraft are temporarily disabled and reverse directions, a rotor deceleration state in which the main rotor assembly is brought to a stop, or a forward flight preparation state in which the at least one wing is rotated to a forward flight position and the main rotor assembly is moved to the aft position so as to position the center of pressure aft of the center of gravity, and wherein the at least one backward transition state includes at least one of a quad hover state in which the at least one wing is rotated to a VTOL flight position and the main rotor assembly is moved to the forward position such that the center of pressure is approximatelyAttorney Docket No.: MIT 25110 PCT | 88212-425635 aligned with the center of gravity, or a rotor acceleration state in which the main rotor assembly is brought to a nominal operating speed. 78. The method of example 76 or 77, wherein the control scheme is further configured to at least one of transition to the kill state from the at least one flight state, from the disarm state, and from the arm state, transition to the disarm state from the kill state, or transition from the disarm state to the arm state. 79. The method of example 77 or 78, wherein the control scheme is further configured to at least one of transition from rotor deceleration state to the arm state or transition from the arm state to the quad hover state and the rotor spin up state. 80. The method of any of examples 76 to 79, wherein the control scheme is further configured to at least one of transition from the quad hover state to the forward flight prep state and the rotor acceleration state, transition from the forward flight prep state to the forward flight state, transition from the rotor acceleration state to the VTOL state, transition from the rotor spin up state to the VTOL state, transition from the VTOL state to the deceleration prep state, transition from the deceleration prep state to the rotor deceleration state, transition from the rotor deceleration state to the forward flight prep state, or transition from the forward flight state to the forward flight prep state. 81. An aircraft, comprising: a base frame assembly; a main rotor assembly including at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, the at least one wing defining a center of pressure; and at least one component adjustment assembly arranged on the base frame assembly and configured to one or more of move the center of pressure of the at least one wing relative to a center of gravity of the aircraft or move the center of gravity relative to the center of pressure. 82. The aircraft of example 81, wherein the at least one component adjustment assembly is configured to move at least one component of the aircraft relative to the base frame assembly to one or more of move the center of pressure of the at least one wing relative to a center of gravity of the aircraft or move the center of gravity relative to the center of pressure.Attorney Docket No.: MIT 25110 PCT | 88212-425635 83. The aircraft of example 82, wherein the at least one component of the aircraft that is movable by the at least one component adjustment assembly includes the main rotor assembly. 84. The aircraft of example 83, wherein the at least one component adjustment assembly is a first component adjustment assembly and is configured to move the main rotor assembly in a forward-aft direction to move the center of pressure of the at least one wing in the forward-aft direction. 85. The aircraft of any of examples 81 to 84, wherein the first component adjustment assembly includes a first carriage configured to move in the forward-aft direction relative to the base frame assembly, and wherein the main rotor assembly is coupled to the first carriage of the first component adjustment assembly such that movement of the first carriage relative to the base frame assembly moves the center of pressure of the at least one wing forward and aft relative to the center of gravity of the aircraft. 86. The aircraft of example 85, wherein the first carriage is movable between a forward position for VTOL flight of the aircraft in which the center of pressure of the at least one wing is approximately aligned with the center of gravity of the aircraft in the forward-aft direction and an aft position for forward flight of the aircraft in which the center of pressure of the at least one wing is aft of the center of gravity of the aircraft in the forward-aft direction. 87. The aircraft of example 85 or 86, wherein the first component adjustment assembly further includes a first linear rail arranged on an upper side of the base frame assembly that extends in the forward-aft direction, and wherein the first carriage is slidably movable on the first linear rail. 88. The aircraft of any of examples 81 to 87, wherein the at least one component of the aircraft that is movable by the component adjustment assembly further includes a first component of the aircraft, wherein the at least one component adjustment assembly further includes a second component adjustment assembly, andAttorney Docket No.: MIT 25110 PCT | 88212-425635 wherein the second component adjustment assembly is configured to move the first component in a forward-aft direction to move the center of gravity of the aircraft in the forward-aft direction. 89. The aircraft of example 88, wherein the first component is an electronics component of the aircraft. 90. The aircraft of example 88 or 89, wherein the second component adjustment assembly includes a second carriage configured to move in the forward-aft direction relative to the base frame assembly, and wherein the first component is coupled to the second carriage of the second component adjustment assembly such that movement of the second carriage relative to the base frame assembly moves the center of gravity forward and aft relative to the center of pressure of the at least one wing. 91. The aircraft of example 90, wherein the second carriage is movable between an aft position for VTOL flight of the aircraft in which the center of pressure of the at least one wing is approximately aligned with the center of gravity of the aircraft in the forward-aft direction and a forward position for forward flight of the aircraft in which the center of pressure of the at least one wing is aft of the center of gravity of the aircraft in the forward-aft direction. 92. The aircraft of example 90 or 91, wherein the second component adjustment assembly further includes a second linear rail arranged on an underside of the base frame assembly that extends in the forward-aft direction, and wherein the second carriage is slidably movable on the second linear rail.

[0339] It should be understood that the subject matter defined in the appended claims is not necessarily limited to the specific implementations described above. The specific implementations described above are disclosed as examples only. One skilled in the art will appreciate further features and advantages of the disclosure based on the above-described embodiments. Accordingly, the disclosure is not to be limited by what has been particularly shown and described, except as indicated by the appended claims. For example, while the present embodiments often include a single feature, it is possible that multiple of the sameAttorney Docket No.: MIT 25110 PCT | 88212-425635 features can be incorporated into the design of a stop-rotor aircraft similar to the aircraft described above without departing from the spirit of the present disclosure. Any and all combinations of features envisioned by a person skilled in the art based on the present disclosure and knowledge in the art are contemplated by the description herein.

[0340] Some non-limiting claims that are supported by the contents of the present disclosure are provided below.

Claims

1. Attorney Docket No.: MIT 25110 PCT | 88212-425635 What is claimed is:

1. An aircraft, comprising: a base frame assembly; a main rotor assembly including at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, the at least one wing defining a center of pressure; and a component adjustment assembly arranged on the base frame assembly and configured to move at least one component of the aircraft relative to the base frame assembly to one or more of move the center of pressure of the at least one wing relative to a center of gravity of the aircraft or move the center of gravity relative to the center of pressure.

2. The aircraft of claim 1, wherein the at least one component of the aircraft that is movable by the component adjustment assembly includes the main rotor assembly.

3. The aircraft of claim 2, wherein the component adjustment assembly is configured to move the main rotor assembly in a forward-aft direction to move the center of pressure of the at least one wing in the forward-aft direction.

4. The aircraft of claim 3, wherein the component adjustment assembly includes a carriage configured to move in the forward-aft direction relative to the base frame assembly, and wherein the main rotor assembly is coupled to the carriage of the component adjustment assembly such that movement of the carriage relative to the base frame assembly moves the center of pressure of the at least one wing forward and aft relative to the center of gravity of the aircraft.

5. The aircraft of claim 4, wherein the carriage is movable between a forward position for VTOL flight of the aircraft in which the center of pressure of the at least one wing is approximately aligned with the center of gravity of the aircraft in the forward-aft direction and an aft position for forward flight of the aircraft in which the center of pressure of the at least one wing is aft of the center of gravity of the aircraft in the forward-aft direction.Attorney Docket No.: MIT 25110 PCT | 88212-425635 6. The aircraft of claim 5, wherein the component adjustment assembly further includes a linear rail arranged on the base frame assembly that extends in the forward-aft direction, and wherein the carriage is slidably movable on the linear rail.

7. The aircraft of claim 6, wherein the component adjustment assembly further includes a linear actuator and an actuator rod coupled to the carriage, and wherein the linear actuator is configured to move the actuator rod toward and away from the linear actuator to move the carriage toward and away from the linear actuator along the linear rail.

8. The aircraft of claim 1, further comprising: a multicopter assembly arranged on the base frame assembly and including a multicopter plate and a plurality of multicopter rotors arranged on the multicopter plate, the plurality of multicopter rotors configured to generate thrust in a downward direction so as to provide lift to the aircraft.

9. The aircraft of claim 8, wherein the multicopter plate is arranged on the base frame assembly above the center of gravity of the aircraft such that a center of thrust of the plurality of multicopter rotors is arranged above the center of gravity.

10. The aircraft of claim 1, wherein the at least one component of the aircraft that is movable by the component adjustment assembly includes a first component of the aircraft, and wherein the component adjustment assembly is configured to move the first component in a forward-aft direction to move the center of gravity of the aircraft in the forward-aft direction.

11. A method of operating an aircraft, comprising: operating an aircraft in a VTOL configuration in which at least one component of the aircraft is arranged relative to the base frame assembly such that a center of pressure of at least one wing of a main rotor assembly of the aircraft is approximately aligned with a center of gravity of the aircraft in a forward-aft direction; moving the at least one component relative to the base frame assembly to a forwardAttorney Docket No.: MIT 25110 PCT | 88212-425635 flight configuration such that the center of pressure of the at least one wing is aft of the center of gravity of the aircraft; and operating the aircraft in a forward flight configuration.

12. The method of claim 11, wherein the aircraft includes a base frame assembly, the main rotor assembly including the at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, the at least one wing defining the center of pressure, and at least one component adjustment assembly arranged on the base frame assembly and configured to move at least one component of the aircraft relative to the base frame assembly.

13. The method of claim 12, wherein the at least one component of the aircraft that is movable by the at least one component adjustment assembly includes the main rotor assembly, the method further comprising: moving the main rotor assembly via the at least one component adjustment assembly in a forward-aft direction to move the center of pressure of the at least one wing in the forward-aft direction.

14. The method of claim 13, further comprising: moving a carriage of the at least one component adjustment assembly to which the main rotor assembly is coupled in the forward-aft direction relative to the base frame assembly such that movement of the carriage relative to the base frame assembly moves the center of pressure of the at least one wing forward and aft relative to the center of gravity of the aircraft.

15. The method of claim 14, further comprising: sliding the carriage along a linear rail of the at least one component adjustment assembly arranged on the base frame assembly.

16. The method of claim 12, wherein the at least one component of the aircraft that is movable by the at least one component adjustment assembly includes a first component of the aircraft, the method further comprising: moving the first component via the at least one component adjustment assembly in a forward-aft direction to move the center of gravity of aircraft in the forward-aft direction.Attorney Docket No.: MIT 25110 PCT | 88212-425635 17. An aircraft, comprising: a base frame assembly; a main rotor assembly including at least one wing extending away from and configured to rotate about an axis such that the at least one wing rotates about the axis to produce lift, the at least one wing defining a center of pressure; and at least one component adjustment assembly arranged on the base frame assembly and configured to one or more of move the center of pressure of the at least one wing relative to a center of gravity of the aircraft or move the center of gravity relative to the center of pressure.

18. The aircraft of claim 17, wherein the at least one component adjustment assembly is configured to move at least one component of the aircraft relative to the base frame assembly to one or more of move the center of pressure of the at least one wing relative to a center of gravity of the aircraft or move the center of gravity relative to the center of pressure.

19. The aircraft of claim 18, wherein the at least one component adjustment assembly is a first component adjustment assembly and is configured to move the main rotor assembly in a forward-aft direction to move the center of pressure of the at least one wing in the forward-aft direction.

20. The aircraft of claim 19, wherein the at least one component of the aircraft that is movable by the component adjustment assembly further includes a first component of the aircraft, wherein the at least one component adjustment assembly further includes a second component adjustment assembly, and wherein the second component adjustment assembly is configured to move the first component in a forward-aft direction to move the center of gravity of the aircraft in the forward-aft direction.

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