Tilt device, aircraft and method

Abstract

A system for an aircraft is provided, the aircraft including a tiltable prop-rotor tiltable between a lift position for providing lift to the aircraft and a forward flight position for providing forward thrust to the aircraft, the system including at least one actuator for adjusting a tilt angle of the tiltable prop-rotor, and at least one passive damper connected to the tiltable prop-rotor and configured to limit a rate of change of the tilt angle of the tiltable prop-rotor.

Description

[Technical Field]

[0001] CROSS-REFERENCE TO RELATED APPLICATIONS This application claims the benefit of U.S. Provisional Application No. 63 / 378,536, filed October 6, 2022, and U.S. Provisional Application No. 63 / 378,680, filed October 7, 2022, and U.S. Application No. 18 / 296,062, filed April 5, 2023, the entire contents of each of which are incorporated herein by reference.

[0002] The present disclosure relates generally to aircraft, and more particularly to aircraft having tiltable propellers. [Background technology]

[0003] A vertical take-off and landing (VTOL) aircraft is an aircraft capable of vertical take-off, landing, and hovering. To do so, a VTOL aircraft can include one or more propellers that can tilt between a position to provide vertical thrust for take-off, landing, and hovering and a position to provide forward thrust for forward flight. A VTOL aircraft can include wings, like a conventional fixed-wing aircraft, that provide lift during forward flight to which tiltable propellers can be attached. Because tiltable propellers control the performance of the aircraft in flight, they must be designed to avoid catastrophic failure. This often involves introducing redundant actuators to control the tilt of the propellers so that if one actuator fails, another actuator takes over, and / or over-designing aircraft structural components so that certain failure modes are not catastrophic. However, redundant actuators and over-designed components introduce additional and undesirable weight that can increase the manufacturing and operating costs of the aircraft. Summary of the Invention

[0004] Disclosed herein is a secondary load path comprising a passive damper for an actuated tiltable prop rotor of an aircraft configured to damp high-speed tilting of the tiltable prop rotor if the actuator tilting the tiltable prop rotor becomes disconnected from the tiltable prop rotor. The shaft of the actuator used to tilt the tiltable prop rotor can be a single point of failure that can experience relatively high loads. Shaft failure can result in uncontrolled tilting of the prop rotor, which can lead to high rotational speeds and high support structure impact velocities under load, resulting in catastrophic failure. The passive damper described herein is configured to damp the rotational speed of the tiltable component in the event of actuator disconnection, thus preventing catastrophic failure without requiring redundancy and over-design of the support structure.

[0005] In some examples, a damper configured to damp high-speed tilting of a tiltable prop-rotor as described herein may include a balanced hydraulic or pneumatic cylinder (a "balanced unfilled damper"). The balanced unfilled damper cylinder may include a piston that slides within the damper's cylinder, the piston including a plurality of openings through which damping fluid flows while the piston moves through the cylinder. A piston rod connected to the damper's piston may be rotatably connected to the tiltable prop-rotor such that the damper applies a force in a direction opposite to the rotation of the tiltable prop-rotor when the prop-rotor tilts toward either a climb configuration or a forward flight configuration. This force is a result of the resistance of the damping fluid applied to the damper piston as it moves through the fluid. Thus, the damper dissipates energy (e.g., by converting kinetic energy to heat) as the piston moves through the damping fluid due to the resistance of the damping oil applied to the piston. In some examples, the damper may include an unbalanced cylinder (an "unbalanced damper"). An unbalanced damper similarly includes a piston that slides within a cylinder and includes a plurality of openings through which damping fluid flows while the piston moves through the cylinder, but may also include pressurized gas within the cylinder in addition to the damping fluid. When a force is applied to the damper piston, the pressurized gas can compress or expand depending on the direction of the force. In this way, the pressurized gas can absorb the initial, sudden force applied to the damper piston.

[0006] According to certain aspects, a tiltable prop-rotor of an aircraft tiltable between a lift position and a forward flight position is connected to a secondary load path including a passive damper, such as the balanced unloaded damper or unbalanced damper described above. In the lift position, the tiltable prop-rotor is configured to provide lift to the aircraft, for example, during vertical takeoff. In the forward flight position, the tiltable prop-rotor is configured to provide forward thrust to the aircraft. The aircraft may include an actuator connected to the tiltable prop-rotor to adjust the tilt angle of the tiltable prop-rotor between the lift position and the forward flight position. As described above, high loads experienced by the actuator shaft may cause the shaft to break, thus resulting in a free-floating tiltable prop-rotor, which may cause catastrophic failure of the aircraft. Thus, the aircraft disclosed herein includes at least one passive damper connected to the tiltable prop-rotor and configured to limit the rate of change of the tilt angle of the tiltable prop-rotor.

[0007] An exemplary aircraft is provided that includes a tiltable prop-rotor tiltable between a climb position to provide lift to the aircraft and a forward flight position to provide forward thrust to the aircraft, at least one actuator to adjust the tilt angle of the tiltable prop-rotor, and at least one passive damper connected to the tiltable prop-rotor and configured to limit a rate of change of the tilt angle of the tiltable prop-rotor.

[0008] In some embodiments of the aircraft, the at least one passive damper includes a hydraulic or pneumatic cylinder.

[0009] In some aircraft embodiments, the hydraulic or pneumatic cylinder is a balanced non-filled cylinder or an unbalanced cylinder.

[0010] In some embodiments of the aircraft, the at least one passive damper is configured to limit the rate of change of the tilt angle of the tiltable prop-rotor in both tilt directions.

[0011] In some embodiments of the aircraft, the at least one actuator includes a linear actuator.

[0012] In some embodiments of the aircraft, the at least one actuator includes a rotary actuator.

[0013] In some embodiments of the aircraft, the aircraft includes a boom and the tiltable prop rotor is tiltably mounted to the boom.

[0014] In some embodiments of the aircraft, the aircraft includes a fixed rotor mounted on a boom at a fixed location to provide lift.

[0015] In some embodiments of the aircraft, the tiltable prop rotor is mounted at the forward end of the boom and the fixed rotor is mounted at the aft end of the boom.

[0016] In some embodiments of the aircraft, the boom includes a hull, and the at least one passive damper is at least partially housed within the hull.

[0017] In some embodiments of the aircraft, the boom comprises a rib, and the at least one actuator and the at least one passive damper are positioned on opposite sides of the rib.

[0018] In some aircraft embodiments, the boom is mounted to the aircraft wing, inside the tip of the wing.

[0019] In some embodiments of the aircraft, the tiltable prop rotor is configured to tilt in an upward direction from a forward flight position to a climb position, and a force vector of the at least one passive damper extends below a tilt axis of the tiltable prop rotor.

[0020] In some embodiments of the aircraft, the aircraft includes a plurality of tiltable prop-rotors and a plurality of passive dampers for the plurality of tiltable prop-rotors.

[0021] In some embodiments of the aircraft, the tilt angle range of the tiltable prop rotor is at least 90 degrees.

[0022] In some embodiments of the aircraft, the aircraft includes a single actuator for adjusting the tilt angle of the tiltable propeller.

[0023] In some embodiments of the aircraft, the at least one passive damper is configured to limit a rate of change of the tilt angle to a predetermined threshold when the tiltable prop-rotor is disconnected from the at least one actuator during forward flight.

[0024] In some embodiments of the aircraft, the tiltable propeller is electrically driven.

[0025] In some embodiments of the aircraft, the aircraft is a passenger airliner.

[0026] An exemplary method for controlling an aircraft includes receiving, at a controller, a command to adjust a tilt angle of a tiltable propeller that is tiltable between a lift position to provide lift to the aircraft and a forward flight position to provide forward thrust to the aircraft, and controlling at least one actuator to adjust the tilt angle of the tiltable propeller according to the command, wherein at least one passive damper is coupled to the tiltable propeller to limit a rate of change of the tilt angle of the tiltable propeller. This method may be performed using any of the aircraft described above.

[0027] It will be understood that any of the variations, aspects, features, and options described in terms of the system apply equally to the method, and vice versa. It will also be apparent that any one or more of the variations, aspects, features, and options described above may be combined. [Brief explanation of the drawings]

[0028] The invention will now be described, by way of example only, with reference to the accompanying drawings, in which:

[0029] [Figure 1] 1 illustrates an aircraft in a forward flight configuration, in accordance with one or more embodiments of the present disclosure.

[0030] [Figure 2] 2 is a perspective view of the aircraft of FIG. 1 illustrating prop rotor positions in a climb configuration and a forward flight configuration in accordance with one or more embodiments of the present disclosure.

[0031] [Figure 3] 1 illustrates a side view of a portion of an aircraft having an exemplary system, in accordance with one or more embodiments of the present disclosure.

[0032] [Figure 4] 1 illustrates a detailed diagram of an exemplary system in accordance with one or more embodiments of the present disclosure.

[0033] [Figure 5] 5 illustrates a detailed view of a portion of the example system of FIG. 4 in accordance with one or more embodiments of the present disclosure.

[0034] [Figure 6] 5 illustrates another detailed view of a portion of the example system of FIG. 4, in accordance with one or more embodiments of the present disclosure.

[0035] [Figure 7] 1A and 1B show a top perspective view and a detailed view of an exemplary system according to one or more embodiments of the present disclosure.

[0036] [Figure 8] FIG. 1 illustrates a detailed view of an exemplary system having a tensioned control rod, in accordance with one or more embodiments of the present disclosure.

[0037] [Figure 9] 9 illustrates a detailed cutaway view of a bushing and cam interface of the example system of FIG. 8 in accordance with one or more embodiments of the present disclosure.

[0038] [Figure 10] 1 illustrates a cutaway detail view of an exemplary hub of a prop rotor, in accordance with one or more embodiments of the present disclosure.

[0039] [Figure 11] 1 shows an example graph of blade pitch as a function of tilt angle for a prop rotor.

[0040] [Figure 12] FIG. 4 is a diagram illustrating an example of a blade pitch angle.

[0041] [Figure 13] 1 illustrates a detailed diagram of an exemplary system in accordance with one or more embodiments of the present disclosure.

[0042] [Figure 14]1 illustrates a detailed diagram of an exemplary system in accordance with one or more embodiments of the present disclosure.

[0043] [Figure 15A] FIG. 2 illustrates a detailed view of an exemplary tiltable prop rotor damping system in a first configuration, in accordance with one or more embodiments of the present disclosure.

[0044] [Figure 15B] FIG. 1 illustrates a detailed view of an exemplary tiltable prop rotor damping system in a second configuration, in accordance with one or more embodiments of the present disclosure.

[0045] [Figure 16] FIG. 2 illustrates a detailed view of an exemplary tiltable prop rotor damping system, in accordance with one or more embodiments of the present disclosure.

[0046] [Figure 17] 10 shows an example graph of damping force as a function of linear velocity of a damper piston, in accordance with one or more embodiments of the present disclosure.

[0047] [Figure 18A] 10 illustrates a graph of prop rotor and damped hinge moment as a function of time in accordance with one or more embodiments of the present disclosure.

[0048] [Figure 18B] 1 illustrates a graph of prop rotor angular velocity as a function of time in accordance with one or more embodiments of the present disclosure.

[0049] [Figure 18C] 10 shows a graph of the linear velocity of a damper piston as a function of time in accordance with one or more embodiments of the present disclosure.

[0050] [Figure 18D] 10 shows a graph of damper force as a function of time in accordance with one or more embodiments of the present disclosure.

[0051] [Figure 19] 1 shows a graph of an example damper effective moment arm length as a function of prop rotor tilt angle for a first damper configuration, in accordance with one or more embodiments of the present disclosure.

[0052] [Figure 20] 10 shows a graph of an example damper effective moment arm length as a function of prop rotor tilt angle for a second damper configuration, in accordance with one or more embodiments of the present disclosure.

[0053] [Figure 21] 1 illustrates a side view of an exemplary system in a first configuration, in accordance with one or more embodiments of the present disclosure.

[0054] [Figure 22] 16 illustrates the example system of FIG. 15 in a second configuration, in accordance with one or more embodiments of the present disclosure.

[0055] [Figure 23] 16 illustrates a front view of the example system of FIG. 15 in accordance with one or more embodiments of the present disclosure.

[0056] [Figure 24A] 1 illustrates an exemplary ratchet configuration. [Figure 24B] 1 illustrates an exemplary locking arrangement for locking a prop rotor in at least one direction.

[0057] [Figure 25] 1 illustrates a front view of an exemplary system according to one or more embodiments of the present disclosure. DETAILED DESCRIPTION OF THE INVENTION

[0058] In the following description of various embodiments, reference is made to the accompanying drawings which show, by way of illustration, specific embodiments which can be practiced. The description is presented to enable any person skilled in the art to make and use the invention and is provided in the context of a patent application and its requirements. Various modifications to the described embodiments will be readily apparent to those skilled in the art, and the generic principles herein may be applied to other embodiments. Thus, the present invention is not intended to be limited to the embodiments shown but is to be accorded the widest scope consistent with the principles and features described herein.

[0059] Described herein are systems, devices, and methods for damping high-speed tilting of an aircraft tiltable prop rotor using a passive damper connected between the aircraft tiltable prop rotor and a support structure. Uncontrolled movement of the tiltable prop rotor poses a risk of catastrophic failure if the actuator becomes disconnected (e.g., if a break in the actuator shaft disconnects a connected portion of the actuator from the remainder of the actuator). Actuator shaft failure can result in uncontrolled tilting of the tiltable prop rotor, which, under load, can lead to high rotational speeds and therefore high support structure impact velocities, resulting in catastrophic failure of the aircraft.

[0060] To mitigate such risks, aircraft may be designed with redundant components, such as redundant actuators, and / or over-designed components that can withstand uncontrolled movement, both of which add weight and cost to the aircraft design and increase operating costs. In accordance with the principles herein, a secondary load path including a passive damper, as described further below, can eliminate the need for redundant actuators and / or over-design of aircraft structural components to mitigate uncontrolled movement of the tiltable prop rotor by damping high speed tilting of the tiltable prop rotor when the actuator controlling the tilting of the tiltable prop rotor is disconnected from the tiltable prop rotor.

[0061] A tiltable propeller of an aircraft tiltable between a lift position and a forward flight position can be connected to a passive damper as described above according to various embodiments. In the lift position, the tiltable propeller is configured to provide lift to the aircraft, for example, during takeoff. In the forward flight position, the tiltable propeller is configured to provide forward thrust to the aircraft. To adjust the angle of the tiltable propeller (e.g., between the lift and forward flight positions), the aircraft can include an actuator between each tiltable propeller and an aircraft structure, such as a boom. Due to the high loads experienced by the actuator shaft, a failure mode that must be considered in the aircraft design is actuator shaft breakage, which can result in a free-floating component (tilt propeller), which can cause catastrophic failure of the aircraft due to high-speed impact of the tilt propeller against an aircraft structural component. Therefore, the aircraft disclosed herein includes at least one passive damper connected to each tiltable propeller and configured to limit the rate of change of the tilt angle of the tiltable propeller.

[0062] The passive damper described herein may include a balanced hydraulic or pneumatic cylinder (a "balanced non-filling damper"). The balanced non-filling damper cylinder may include a piston that slides within the cylinder and includes a plurality of openings through which damping fluid flows while the piston moves through the cylinder. The piston may be attached to a piston rod that extends from the cylinder and may be rotatably attached to a tiltable propeller of an aircraft. As the tiltable propeller rotates toward either a raised position or a forward flight position, it exerts a tensile or compressive force on the piston rod, respectively, moving the piston rod away from or toward the piston cylinder. Thus, as the propeller tilts in either direction, the piston is forced to slide within the cylinder. In this manner, the damper dissipates energy (e.g., by converting kinetic energy to heat) as the piston moves through the damping fluid due to the resistance of the damping fluid applied to the piston.

[0063] In some examples, the passive damper may include an unbalanced cylinder (“unbalanced damper”). The unbalanced damper similarly includes a piston that slides within the cylinder and includes multiple openings through which damping fluid flows while the piston moves through the cylinder, but may also include pressurized gas within the cylinder in addition to the damping fluid. When a force is exerted on the damper piston, the pressurized gas can compress or expand depending on the direction of the force. In this way, the pressurized gas can absorb the initial, sudden force exerted on the damper piston. Thus, in the event of actuator shaft failure, the passive damper (either a balanced, uncharged damper or an unbalanced damper) damps the high-speed tilting of the tiltable prop rotor, thereby preventing catastrophic failure.

[0064] The systems, devices, and methods described herein therefore provide for more cost-effective aircraft designs and overall aircraft weight reduction while meeting safety requirements. Furthermore, passive dampers do not require actuation, engagement, or other control to perform their function of controlling the rate of change of tilt of a tiltable component, thus providing a reliable and simple solution to the actuator disconnect failure mode described herein.

[0065] As used herein, the singular forms "a," "an," and "the" used in the following description are intended to include the plural forms as well, unless the context clearly dictates otherwise. The term "and / or," as used herein, will also be understood to refer to and include any and all possible combinations of one or more of the associated listed items. It will be further understood that the terms "includes," "including," "comprise," and / or "comprising," as used herein, specify the presence of stated features, integers, steps, operations, elements, and / or components, but do not exclude the presence or addition of one or more other features, integers, steps, operations, elements, components, and / or groups thereof.

[0066] As used herein, the term "prop rotor" refers to a variable tilt propeller that can change the direction of propeller thrust by changing the tilt angle of the propeller. For example, the tilt angle can be changed from an angle that provides at least some vertical thrust, such as for vertical takeoff and landing, to an angle that provides at least some horizontal thrust, such as for forward flight. As used herein, a propeller climb configuration refers to any tilt propeller orientation in which the propeller thrust is primarily providing lift to the aircraft, and a propeller forward flight configuration refers to any propeller orientation in which the propeller thrust is primarily providing forward thrust to the aircraft.

[0067] As used herein, "vertical takeoff and landing" ("VTOL") refers to the ability of an aircraft to move substantially vertically without lift provided solely by the aircraft's wings. The term directly encompasses vertical takeoff and landing (i.e., vertical movement without horizontal movement), but also vertical movement in combination with horizontal movement. Those skilled in the art will understand that a VTOL aircraft may be capable of non-vertical takeoff and landing. For example, a winged VTOL, such as the various embodiments described herein, can take off and land in the manner of a conventional airplane, utilizing the lift provided by its wings at an appropriate airspeed.

[0068] 1 shows an aircraft 100 in a forward flight configuration. The aircraft 100 includes a fuselage 102, wings 104 attached to the fuselage 102, and one or more rear stabilizers 106 attached to the rear of the fuselage 102. The aircraft 100 may be a vertical take-off and landing (VTOL) aircraft and may be a passenger airliner. A plurality of rotors 112 are attached to the wings 104 and configured to provide lift for take-off, landing, and the like. A plurality of propellers 114 are attached to the wings 104 and are tiltable between a climb configuration that provides a portion of the lift required for vertical take-off, landing, and hovering, and a forward flight configuration (as shown in FIG. 1 ) that provides forward thrust to the aircraft 100 for horizontal flight.

[0069] During takeoff and landing, the prop rotors 114 are tilted to a climb configuration in which their thrust is directed upward to provide lift. For forward flight, the prop rotors 114 tilt from their climb configuration to their forward flight configuration in which their thrust is directed forward to provide forward propulsion. In other words, the pitch of the prop rotors 114 varies from a tilt angle at which the prop rotors provide lift for takeoff and landing (and, optionally, hovering) to a tilt angle at which the prop rotors provide forward thrust to the aircraft 100 for forward flight. The prop rotors 114 can each be tilted by one or more actuators. The actuator(s) can be electrically driven. Optionally, each prop rotor has a single actuator for adjusting its tilt. According to various embodiments, the aircraft 100 can include one or more damper mechanisms connected to each tiltable prop rotor configured to limit the rate of change of the tilt angle of the tiltable prop rotor, such as if the actuator is disconnected or otherwise fails.

[0070] When the aircraft 100 is in full forward flight, lift may be provided entirely by the wings 104 and the rotors 112 may be shut off. The blades 120 of the rotors 112 may be locked in a low-drag position for aircraft cruise. In some embodiments, the rotors 112 each have two blades 120 that are locked for cruise in a minimum-drag position with one blade immediately ahead of the other, as illustrated in FIG. 1 . In some embodiments, the rotors 112 have three or more blades. In some embodiments, the prop rotors 114 include more blades 116 than the rotors 112. For example, as illustrated in FIG. 1 , the rotors 112 may each include two blades, and the prop rotors 114 may each include five blades. According to various embodiments, the prop rotors 114 may have between two and five blades.

[0071] According to various embodiments, the aircraft includes only one wing 104 on each side of the fuselage 102 (or a single wing extending across the entire aircraft), with at least a portion of the rotors 112 located aft of the wing 104 and at least a portion of the prop rotors 114 located forward of the wing 104. In some embodiments, all of the rotors 112 are located aft of the wing 104 and all of the prop rotors are located forward of the wing 104. According to some embodiments, all of the rotors 112 and the prop rotors 114 are mounted to the wing, i.e., neither the rotors nor the prop rotors are mounted to the fuselage. According to various embodiments, all of the rotors 112 are located aft of the wing 104 and all of the prop rotors 114 are located forward of the wing 104. According to some embodiments, all of the rotors 112 and the prop rotors 114 are positioned inboard of the wing tips 109.

[0072] According to various embodiments, the rotors 112 and the prop rotors 114 are attached to the wings 104 by booms 122. The booms 122 may be attached below the wings 104, above the wings, and / or integrated into the wing profile. According to various embodiments, one rotor 112 and one prop rotor 114 are attached to each boom 122. The rotors 112 may be attached to the aft ends of the booms 122, and the prop rotors 114 may be attached to the forward ends of the booms 122. In some embodiments, the rotors 112 are attached in fixed positions on the booms 122. In some embodiments, the prop rotors 114 are attached to the forward ends of the booms 122 via hinges or other systems. The prop rotors 114 may be attached to the booms 122 such that when in their forward flight configuration, the prop rotors 114 are aligned with the body of the booms 122 and form a continuous extension of the forward ends of the booms 122, minimizing drag for forward flight.

[0073] During takeoff and landing, an aircraft is operated by positioning the prop-rotors in a climb configuration and providing the aircraft with the necessary lift through the rotors and the combined lift provided by the prop-rotors. According to various embodiments, during takeoff, landing, and / or hovering, the prop-rotors can be maintained in a predetermined climb configuration, which can be the same across all prop-rotors or can be different for different prop-rotors. According to various embodiments, the tilt of at least some of the prop-rotors can be actively adjusted during takeoff, landing, and / or hovering to provide the necessary stability and / or control. As discussed further below, the pitch of the prop-rotor blades (also referred to herein as pitch angle and angle of attack) is mechanically linked to the prop-rotor tilt, such that the blade pitch is coordinated with the prop-rotor tilt, such as to achieve maximum thrust when the prop-rotor is in a climb configuration and improved efficiency when the prop-rotor is in a forward flight configuration.

[0074] According to various embodiments, each rotor and / or each prop-rotor can be individually controlled by a flight controller according to various degrees of freedom of movement. According to various embodiments, the only degree of freedom of a rotor is the rotational speed of the rotor. According to various embodiments, the degrees of freedom of at least a portion of the prop-rotor include the rotational speed of the prop-rotor and the tilt degree of the prop-rotor (combined with the blade pitch of the prop-rotor). According to various embodiments, any of these degrees of freedom can be actively controlled by the flight controller (autonomously or in response to pilot commands) during takeoff and landing to provide appropriate stability and control.

[0075] When the aircraft reaches an altitude sufficient to begin forward flight, the prop rotors begin to tilt forward toward their forward flight configuration so that their thrust provides a combination of lift and thrust, with the rate of climb decreasing as the prop rotors tilt further toward their forward flight configuration. The blade pitch angle can be adjusted as the prop rotors tilt forward toward their forward flight configuration. For example, in the forward flight configuration, the prop rotor blades can be pitched at an angle that provides less drag relative to the blade pitch angle when in a climb configuration. The rotors can remain active for at least a portion of the time the prop rotors are tilted forward and continue to provide rotor-based lift. The rotors can be deactivated any time after the forward airspeed reaches a height sufficient for the wings to provide enough lift to maintain the aircraft's altitude.

[0076] The tilt of at least a portion of the proprotors can be actively controlled to provide additional stability and / or maneuverability control during cruise. In some embodiments, the tilt of at least some of the proprotors is actively controlled during takeoff, landing, and / or hovering. In some embodiments, the tilt of the proprotors is fixed (i.e., does not vary) during cruise. According to some embodiments, the tilt of the outermost proprotor can be actively and independently controlled during vertical takeoff, landing, and / or hovering to provide yawing moment as needed. The tilt angle range of the tiltable proprotor(s) is at least 90 degrees so that the proprotors can tilt between forward flight and climb configurations.

[0077] 2 is a perspective view of the aircraft 100 of FIG. 1 illustrating prop rotor positions in a climb configuration and a forward flight configuration in accordance with one or more embodiments of the present disclosure. The prop rotor 114 can tilt about a tilt axis 118 that is perpendicular to the forward direction of the aircraft. For forward flight, the prop rotor tilts from a climb configuration that provides vertical thrust to a forward flight configuration that provides forward thrust.

[0078] As described further below, the aircraft 100 may include a system for tilting the prop rotor 114 between a climb configuration and a forward flight configuration. The system may mechanically link adjustments to the pitch angle of the blades of the prop rotor 114 to adjustments to the tilt of the prop rotor 114.

[0079] 3 illustrates an example system 202 for coupling prop rotor tilt and prop rotor blade pitch angle adjustment. System 202 is configured to enable prop rotor blade pitch angle to correspond to prop rotor tilt position without requiring separate systems for tilting the prop rotor and adjusting the blade pitch angle. System 202 therefore reduces aircraft complexity and cost.

[0080] The system 202 rotatably couples the prop rotor 203 to a boom 205 of the aircraft and the prop rotor 114 to a boom 122 of the aircraft 100. The system 202 may include a fixed frame 220 for mounting to the aircraft (e.g., the prop rotor 203 may be connected to the boom 205 via a bracket 236 of the fixed frame 220) and a prop rotor frame 222 to which the prop rotor 203 is mounted. The prop rotor frame 222 may be tiltably connected to the fixed frame 220 at a joint 212.

[0081] The system 202 may include one or more arms 204 connected to a prop rotor frame 222. A linear actuator 206 may be connected to the arm(s) 204 that tilts the prop rotor 203 about a joint 213. The actuator 206 may be, for example, a ball screw actuator or a pneumatic actuator. Alternatively, a rotary actuator such as a stepper motor or servo motor may be attached to the joint or may drive a gear train having an output gear located at or engaging a gear located at the joint.

[0082] The system 202 may include a cam 214 that can rotate in response to tilting the prop rotor 203. A control rod (discussed further below) may be operatively coupled to the cam 214 such that the control rod translates as the cam 214 rotates. The control rod is coupled at its opposite end to the blades 250 of the prop rotor 203 such that translation of the control rod adjusts the pitch angle of the blades 250 of the prop rotor 203. Thus, the system 202 couples the pitch angle of the blades 250 of the prop rotor 203 to the tilt of the prop rotor 203. In operation, the aircraft's control system may send prop rotor tilt adjustment commands to the actuator 206. The actuator 206 may be extended or retracted, causing the prop rotor to increase or decrease its degree of tilt. As the tilt of the prop rotor changes, the cam 214 rotates. This, in turn, may translate the control rod, adjusting the pitch of the blades 250.

[0083] 4 shows a detailed view of an exemplary system 302 for mechanically linking the tilting of an aircraft propeller between a vertical thrust position and a forward thrust position and the adjustment of the pitch angle of the propeller blades. System 302 can be used for system 202 of FIG. 3. System 302 can include a stationary frame 320 for mounting to a portion of the VTOL aircraft (such as the fuselage, wing, or boom structure) and a propeller frame 322 for mounting a propeller 303 rotatably mounted to stationary frame 320 at a rotation axis 301. In one or more embodiments, system 302 includes a gear 308, a pinion (not shown), a cam 312, a control rod 314, and a pair of arms 304.

[0084] As shown in FIG. 4, gear 308 is located along rotation axis 301. Gear 308 can be fixed in place relative to stationary frame 320. For example, as shown in FIG. 4, gear 308 is connected to an internal pin 318 that attaches to stationary frame 320. One or more shafts 316 can surround internal pin 318, which is clearly shown by the cutaway view of the left shaft 316 in FIG. 4. A set of bearings 317 is located between each shaft 316 and stationary frame 320 such that shaft 316 is rotatably mounted to stationary frame 320.

[0085] The proximal end of the arm 304 may engage with the prop rotor frame 322. The engagement between the arm 304 and the prop rotor frame 322 may be fixedly connected, such as by bolting or welding the arm 304 to the prop rotor frame 322. Optionally, both the arm 304 and the prop rotor frame 322 may be fixedly connected to the shaft 316. The distal end of the arm 304 may be connected to one or more actuators 306 (see FIG. 5 ) that drive the arm 304 to rotate about the axis of rotation 301. When the actuators 306 drive the arm 304, the prop rotor 303 rotates about the axis of rotation 301.

[0086] The engagement between gear 308, pinion, cam 312, and control rod 314 is more clearly shown in Figure 5, which shows a detailed view of a portion of example system 302 of Figure 4. Pinion 310 is mounted within pinion housing 309, which is coupled to arm 304 and / or prop rotor frame 322 at a fixed location. Pinion 310 is thus rotatably coupled to prop rotor frame 322 such that pinion 310 moves with prop rotor frame 322 (e.g., as prop rotor frame 322 and prop rotor 303 rotate about rotational axis 301). Pinion 310 is also engaged with gear 308. Rotation of pinion housing 309 drives pinion 310 around at least a portion of gear 308, which causes pinion 310 to rotate via toothed engagement with gear 308.

[0087] Cam 312 is fixedly connected to pinion 310, such as via internal pin 311, so as to rotate with pinion 310. Cam 312 is also operably coupled to a first end of control rod 314 so that control rod 314 translates relative to internal pin 311 during at least a portion of a rotation of cam 312. Control rod 314 is coupled at a second end to a blade of the prop rotor (as described below) so that translation of control rod 314 adjusts the pitch angle of the blade.

[0088] When the prop rotor frame 322 rotates about the rotation axis 301 (e.g., to tilt the prop rotor 303), the pinion 310 orbits about the gear 308, which rotates the cam 312 and translates the control rod 314, thereby adjusting the pitch angle of the blades of the prop rotor 303. Thus, the system 302 mechanically links the tilting of the prop rotor 303 and the adjustment of the pitch angle of the blades of the prop rotor 303.

[0089] The control rod 314 may engage the cam 312 via a follower that follows the cam 312 as it rotates. The follower may be, for example, a roller or a pin. As shown in FIG. 5 , the control rod 314 engages the cam 312 via a roller 313. The roller 313 rides along the outer surface of the cam 312 as it rotates. To maintain engagement with the outer surface of the cam 312, the control rod 314 may be compressively biased against the cam 312, such as via a spring (not shown).

[0090] The cam profile (e.g., the shape of the outer surface along which the control rod follows) controls the position of the control rod. The profile of cam 312 can include one or more portions that cause translation of control rod 314 and one or more portions that do not cause translation of control rod 314. For example, cam 312 can have one or more helical portions that cause translation of control rod 314 and / or one or more circular portions that do not cause translation of control rod 314. In the example shown in FIG. 5, cam 312 includes a helical profile that causes translation of control rod 314 continuously throughout the range of tilt of the propeller.

[0091] Figure 6 shows another detailed view of a portion of the example system 302 of Figures 4 and 5. While Figure 5 depicts the control rod 314 engaging two sides of the outer surface of the cam 312 (e.g., in the corner regions of the snail-shaped cam), Figure 6 depicts the control rod 314 engaging only one side of the outer surface of the cam 312. Compared to the position of the control rod in Figure 5, the position of the control rod 314 depicted in Figure 6 can be achieved by rotating the cam 312 in a clockwise direction such that the roller 313 moves in a counterclockwise direction as it follows the surface of the cam 312.

[0092] As the roller 313 moves along the helical portion of the cam 312, the control rod 314 may translate toward or away from the internal pin 311 at the center of the cam 312. For example, if the cam 312 rotates clockwise, the control rod 314 translates away from the internal pin 311 as the roller 313 moves along the helical portion of the cam 312. Conversely, if the cam 312 rotates counterclockwise, the control rod 314 may translate toward the internal pin 311 as the roller 313 moves along the helical portion of the cam 312. As the roller 313 moves along the circular portion of the cam 312, the control rod 314 may remain a constant distance and not translate relative to the internal pin 311. For example, the cam 312 may include a helical profile for the first 210 degrees of rotation and a circular profile for the remaining 150 degrees of rotation such that the control rod 314 translates away from the internal pin 311 only during the first 210 degrees of rotation of the cam 312. As described above, the control rod 314 may be compressively biased against the cam 312 via a spring 315.

[0093] 6 shows an example in which a control rod 314 has a roller 313 that engages the outer surface of a cam 312. Alternative configurations may have different engagements between the control rod and the cam. For example, FIG. 7 shows a top perspective view of an example system 402 including a control rod 414 having a pin 413 that engages with a track 411 of a cam 412. The pin 413 engages with the track 411 of the cam 412 such that as the cam 412 rotates, the cam 412 can push or pull the pin 413 and the control rod.

[0094] System 402 can be used for system 202 of FIG. 3. Similar to the systems discussed above, system 402 connects prop rotor 403 to a portion of an aircraft (such as to boom 401, as shown in FIG. 7). However, unlike the systems discussed above, system 402 includes a control rod 414 having a pin 413 that runs in a track. This configuration more clearly visualizes the detailed pop-out of cam 412, showing pin 413 of control rod 414 engaging track 411 in cam 412. As cam 412 rotates, for example, by rotationally engaging pinion 410 as pinion 410 orbits gear 408, pin 413 can follow track 411, thereby translating control rod 414.

[0095] Gear 408 may be fixed in position relative to a stationary frame 420 that is fixedly mounted to the aircraft. For example, as shown in Figure 7, gear 408 is connected to stationary frame 420. Pinion 410 may be rotatably mounted to stationary frame 420 such that pinion 410 is rotatably coupled to and moves with prop rotor frame 422.

[0096] The track 411 can include a helical portion and a circular portion. As the pin 413 moves along the helical portion of the track 411, the control rod 414 can translate toward or away from the center of the cam 412. However, as the pin 413 moves along the circular section of the track 411, the control rod 414 can remain a fixed distance and not translate relative to the center of the cam 412. Optionally, the track 411 of the cam 412 can have various geometries based on the type of translation desired. To maintain engagement with the track 411, the control rod 414 can be biased in compression or tension against the cam 412, such as via a spring (not shown).

[0097] Another exemplary cam control rod configuration is shown in Figure 8. Unlike system 302, in system 502, control rod 514 is held in tension. Similar to the systems described above, system 502 can include a prop rotor frame 522 rotatably mounted to a stationary frame 520 that is mounted to a portion of the aircraft (such as the fuselage, wing, or boom structure). System 502 can be configured such that one or more actuators (not shown) drive prop rotor frame 522 to rotate about axis of rotation 501 to tilt a prop rotor mounted to prop rotor frame 522 between a vertical thrust position and a forward thrust position.

[0098] Similar to the systems described above, system 502 includes a control rod 514 that engages a cam 512, which rotates based on engagement with a pinion 510 that engages a gear 508. Gear 508 can be fixed in position relative to a stationary frame 520. For example, as shown in FIG. 7 , gear 508 is connected to stationary frame 520. Cam 512 is fixedly connected to pinion 510, such as via shaft 516, so that cam 512 rotates with pinion 510. Pinion 510 can be rotatably mounted to stationary frame 520, such as via a bearing mount to rib 560, so that pinion 510 is rotatably coupled to and moves with prop rotor frame 522. Pinion 510 orbits around gear 508, causing cam 512 to rotate as prop rotor frame 522 moves.

[0099] Control rod 514 includes a clevis 515 and a follower 513, which in this example is a roller. As shown, cam 512 is engaged with follower 513 such that follower 513 rolls along cam 512 as it rotates. Follower 513 is rotatably mounted to clevis 515 on control rod 514. Control rod 514 is in tension (a force is applied to the left of the view in FIG. 8 along axis 514, such as via one or more springs) such that follower 513 is pressed against cam 512. Thus, as follower 513 rolls along cam 512, control rod 514 may translate along axis 505 (depending on the profile of cam 512). As noted above, cam 512 may have any suitable profile to achieve a desired relationship between blade pitch and prop rotor tilt.

[0100] The clevis 515 includes a slot 536 through which the shaft 516 can extend. The clevis 515 is separated from the shaft 516 by a bushing, which is more clearly shown in FIG. 9, which shows a detailed cutaway view of the interface between the bushing 519 and the clevis 515 of the system 502 of FIG. 8. The bushing 519 can include an engagement surface 529 on the area of ​​the bushing 519 that engages with the clevis 515, which is shaped to prevent the clevis 515 from rotating, which in turn prevents a control rod attached to the clevis 515 (e.g., the control rod 514 shown in FIG. 8) from rotating. As shown in FIG. 9, the engagement surface 529 of the bushing 519 is flat, which corresponds to the flat surface of the clevis 515. Optionally, these surfaces can be another shape based on the shape of the corresponding surface of the clevis 515. For example, the engagement between the clevis 515 and the bushing 519 can involve a circular, elliptical, or angled surface.

[0101] As discussed above, translation of a control rod in any of the example systems can adjust the pitch angle of blades on the prop rotor. The control rod can be operatively engaged with multiple blades such that translation of the control rod causes rotation of the blades. FIG. 10 shows a cutaway detail view of an example hub 1000 of a prop rotor, illustrating an example of a coupling between a control rod 1014 and blades 1012 of the prop rotor, in accordance with one or more embodiments of the present disclosure. The hub 1000 can include an engine shaft 1004, a spring 1015, a bearing 1003, a plate 1002, a pitch plate 1005, and multiple links 1006. The blades 1012 are attached to the hub 1000. For example, each blade 1012 can include a blade root 1010 that connects to the link 1006. The control rod 1014 connects to the pitch plate 1005 and / or the plate 1002 of the hub 1000. The other end of the control rod 1014 may be connected to a system configured to translate the control rod 1014 while tilting the prop rotor, as discussed above.

[0102] The engine shaft 1004 connects to a plate 1002, which surrounds a spring 1015 and bearing 1003 and connects to a pitch plate 1005 engaged with a link 1006. The engine shaft 1004 connects to an engine (not shown) at the hub 1000. The plate 1002 is constrained from rotation by a splined interface to the engine shaft 1004 so that the plate 1002 rotates with the engine shaft 1004. The control rod 1014 is prevented from rotating with the plate 1002 via a bearing 1003. The pitch plate 1005 connects to the blade root 1010 via a link 1006. As shown in FIG. 10, the link 1006 is a dogbone link, although other linkage types, such as a pitch link, are contemplated. The spring 1015 can maintain the control rod 1014 in tension or compression. In the embodiment of FIG. 10, a spring 1015 keeps the control rod 1014 in tension by pressing the control rod 1014 against the pitch plate 1005 .

[0103] As the control rod 1014 translates (e.g., axially forward or backward), the plate 1002 and / or pitch plate 1005 translate, which causes the link 1006 to rotate each blade 1012 about its central axis 1008, thereby adjusting the pitch angle of the blades 1012. Rotating each blade 1012 about its central axis 1008 adjusts the pitch angle of the blades 1012. Thus, translating the control rod 1014 can adjust the pitch angle of the blades 1012 of the prop rotor.

[0104] As discussed above, adjusting the pitch of the prop rotor blades based on the prop rotor tilt can allow the propeller's operating characteristics to be tailored to different phases of flight, which can lead to greater efficiency, which can result in less energy demand over the course of flight. The relationship between blade pitch and prop rotor tilt can be selected by selecting a desired cam profile (e.g., the profile of cam 214). Based on the selection of the cam profile, a wide variety of relationships between blade pitch and prop rotor tilt are achievable. Figure 11 is a graph illustrating some examples of these relationships.

[0105] The graph in FIG. 11 shows blade pitch as a function of propeller tilt angle. The propeller tilt angle, provided on the X-axis of the graph in FIG. 11, is the angle of the propeller's rotation axis relative to a line extending parallel to the aircraft's longitudinal axis and intersecting the propeller's rotation axis. The propeller's rotation axis is shown in the example of FIG. 2 as rotation axis 1101-A when the propeller is in a forward flight position (rotation axis 1101-A coincides with a line extending parallel to the aircraft's longitudinal axis 1102 and intersecting the propeller's rotation axis) and rotation axis 1101-B when the propeller is in a climb position. A zero-degree tilt angle on the graph in FIG. 11 corresponds to a propeller having a rotation axis parallel to the aircraft's longitudinal axis 1102 to provide forward thrust, for example, for forward flight, and a 90-degree tilt angle corresponds to a propeller providing vertical thrust, such as for vertical takeoff and landing.

[0106] The blade pitch, provided on the Y-axis of the graph in FIG. 11 , can be defined as the angle between the blade's chord and the plane of rotation and can be measured at a specific point along the blade's length. FIG. 12 illustrates an example of a blade pitch definition. A blade 1208 (only one is shown for simplicity, but it will be understood that each prop rotor includes multiple blades) orbits a prop rotor rotating shaft 1214 and is mounted to rotate about a pitch axis 1206, allowing the pitch 1216 of the blade 1208 to be adjusted. The plane containing the pitch axis 1206 and traversed by the blade 1208 as it rotates can be referred to as the disk plane 1210. The pitch 1216 of the blade 1208 can be defined as the angle between the chord 1204 of the blade 1208 (the line joining the leading and trailing edges of the blade 1208) and a line 1202 lying in the disk plane 1202 that is perpendicular to the pitch axis 1206. It will be understood by those skilled in the art that the pitch values ​​shown in Figure 11 are merely exemplary and that the pitch values ​​will depend on the specific design of the blade and the location along the blade where the pitch is measured. Similarly, it will be understood by those skilled in the art that the prop rotor tilt values ​​shown in Figure 11 are merely exemplary and that different ranges of prop rotor tilt may be used, including negative tilt angles associated with the prop rotor being tilted somewhat downward.

[0107] FIG. 11 includes four different lines 1102-1108 showing four different relationships between blade pitch and propeller tilt. Each line is associated with a different cam profile. Lines 1102, 1104, and 1106 have maximum blade pitch (the maximum for a given line) at zero propeller tilt angle. This may be the blade pitch for forward flight, where a higher blade pitch can provide better efficiency at the relatively high airspeeds of forward flight.

[0108] Line 1108 has a very low blade pitch at zero prop rotor tilt. This can result in the blades producing relatively high drag at high airspeeds, which can be useful for slowing the aircraft, such as during landing. This low blade pitch can also be useful for increasing the efficiency of the prop rotor at low airspeeds, such as during a conventional (airplane-style) takeoff. The blade pitch quickly increases to a maximum, such that the prop rotor needs to be tilted a relatively small amount to achieve maximum blade pitch, which may be the desired blade pitch for forward flight. In this relationship, to achieve a more optimal blade pitch for high speeds in forward flight, the prop rotor needs to be tilted a small amount relative to the minimum tilt (e.g., zero tilt, a small positive tilt, a small negative tilt, etc., as shown).

[0109] Each line 1102-1108 shows blade pitch decreasing to a minimum blade pitch (the minimum for a given relationship) associated with the maximum propeller tilt that can be used for vertical flight and hovering. The minimum blade pitch may be optimal for the low airspeed and high thrust requirements of vertical flight and hovering. Lines 1104 and 1106 illustrate that the minimum blade pitch need not be zero. The particular propeller tilt angle at which the blade pitch minimum is reached can be selected based on the cam profile, as illustrated by the different locations of this point for various relationships (e.g., point 1110).

[0110] Each relationship 1102-1108 is achieved by a different CAM profile. The cam profile can include multiple regions with different shapes to achieve the varying relationships between blade pitch and propeller tilt for the various lines 1102-1108. For example, line 1102 can be achieved by a cam profile including a helical section corresponding to a range of zero-degree propeller tilt relative to a propeller tilt angle at point 1110. The helical section (continuously varying radius) transitions to a circular section (constant radius) that provides a constant blade pitch associated with the propeller tilt angle past point 1110. Line 1106 can be achieved by a first helical section, followed by a second helical section having a different rate of change of radius than the first helical section, followed by a circular section. Line 1104 can be achieved by a cam profile having a continuously varying rate of change of radius.

[0111] The relationship shown in Figure 11 is merely an example showing that many different relationships can be achieved through appropriate selection of cam profiles. Those skilled in the art will understand that any desired relationship between blade pitch and prop rotor tilt can be achieved through appropriate cam profile design.

[0112] Passively damped secondary load path for tiltable components As described above, disconnection of an actuator from a tiltable aircraft component poses the risk of catastrophic failure due to uncontrolled movement of the tiltable component. The passive dampers described herein provide a mechanism for damping high-speed tilting of a tiltable prop rotor in the event of an actuator disconnection. While the likelihood of catastrophic failure due to actuator disconnection can potentially be mitigated by introducing redundant actuators so that another actuator takes over if the first actuator fails (e.g., due to actuator shaft breakage) and / or by over-designing aircraft structural components so that certain failure modes are not catastrophic, these mitigation methods are expensive and can introduce undesirable additional weight into the aircraft design. The passive dampers described herein provide a modest, cost-effective method of preventing catastrophic failure.

[0113] A passive damper is configured to apply a force (hydraulic or pneumatic) to a connected component only when the connected component is moving (i.e., no bias is applied when the connected component is stationary). The passive damper described herein can be connected between an aircraft boom and a tiltable propeller to apply a damping force when the tilt angle of the tiltable propeller is adjusted. As described above, the passive damper can include a balanced hydraulic or pneumatic cylinder (a "balanced non-loaded damper"). The balanced non-loaded damper cylinder can include a piston that slides within the cylinder, the piston including a plurality of openings through which damping fluid flows while the piston moves through the cylinder. The piston can be attached to a piston rod that extends from the cylinder and can be rotatably attached to the tiltable propeller of the aircraft. When the tiltable propeller rotates toward either a raised position or a forward flight position, it exerts a tensile or compressive force on the piston rod, urging the piston rod either away from or toward the piston cylinder, depending on the direction of tilt. Thus, when the prop rotor tilts in either direction, the piston is forced to slide within the cylinder. Due to the resistance of the damping oil applied to the piston, the damper dissipates energy (e.g., by converting kinetic energy to heat) as the piston moves through the damping fluid, thus damping the tilting speed of the tiltable prop rotor.

[0114] In some examples, the passive damper may include an unbalanced cylinder (“unbalanced damper”) instead of a balanced hydraulic or pneumatic cylinder. The unbalanced damper similarly includes a piston sliding within the cylinder and including multiple openings through which damping fluid flows while the piston moves through the cylinder, but may also include pressurized gas within the cylinder in addition to the damping fluid. When a force is applied to the damper piston, the pressurized gas can compress or expand depending on the direction of the force. In this way, the pressurized gas can absorb the initial, sudden force applied to the damper piston. Therefore, in the event of actuator shaft failure, the damper damps the high-speed tilting of the tiltable prop rotor, thus preventing catastrophic failure. Additionally, the passive damper does not require actuation, engagement, or other control to perform its function of damping the high-speed tilting of the tiltable prop rotor. Therefore, it provides a reliable and simple solution to the actuator disconnect failure mode described herein.

[0115] In one or more embodiments, any of the systems described above, such as system 202, system 302, system 402, system 502, and aircraft 100, may include a damper to limit the tilt rate of the prop rotor, such as in the event of an actuator failure, as shown in example system 1302 of Figure 13. System 1302 may be implemented in aircraft having tiltable prop rotors in place of, or in addition to, redundant actuators and / or joint assemblies including failsafe latches, as described below.

[0116] The dampers 1325 may each include a balanced hydraulic or pneumatic cylinder 1316, as known in the art, and include a piston 1318 sliding within the cylinder and connected to a piston rod 1307 or 1407, the piston including a plurality of openings (not shown) through which fluid flows as the piston moves within the cylinder. The dampers 1325 of the system 1302 are connected between a boom 1305 of an aircraft, such as the aircraft 100 of FIGS. 1 and 2 , and a prop rotor frame 1322 to which a prop rotor (not shown) is mounted. The dampers 1325 may be housed within the outer shell of the boom 1305. In one or more embodiments, the boom 1305 may include a rib 1350, and the actuators 1306 and dampers 1325 may be positioned on either side of the rib 1350. The prop rotor frame 1322 is tiltably connected to the stationary frame 1320 such that the prop rotor frame 1322 (and the prop rotor) can tilt about the axis of rotation 1301. Actuator 1306 is connected to the prop rotor frame 1322 and drives the tilting of the prop rotor about axis of rotation 1301 .

[0117] The damper 1325 may be rotatably attached to the boom 1305 by one or more mounts 1314 on a first side to a rib 1350 and on a second side to the hull of the boom 1305. The damper 1325 is rotatably connected to the prop rotor frame 1322 at a connector 1340 of a piston rod 1307 of the damper 1325. The connector 1340 of the piston rod 1307 of the damper 1325 may be rotatably connected to a pin 1308 that extends between two sides of a clevis bracket 1309. Thus, as the actuator 1306 tilts the prop rotor frame 1322 toward the raised position (clockwise relative to the view of FIG. 13 ), the prop rotor frame 1322 exerts a pulling force on the piston rod 1307 of the damper 1325, urging the piston rod 1307 to extend from its cylinder. Fluid within the cylinder resists such motion, providing a countervailing force acting against the piston 1318 in the damper 1325 that is proportional to the rate of piston movement. Multiple openings in the piston allow damping fluid to flow through the piston 1314 so that the piston can move within the cylinder, thereby allowing the piston rod 1307 to extend from the cylinder until the prop rotor reaches a raised position. Similarly, as the actuator 1306 tilts the prop rotor frame 1322 toward a forward flight position (counterclockwise relative to the view of FIG. 13 ), the prop rotor frame 1322 exerts a compressive force on the piston rod 1307 of the damper 1325, urging the piston rod 1307 to retract into the cylinder 1316. Fluid within the cylinder similarly resists such motion, providing a countervailing force acting against the piston 1318 in the damper 1325 that is proportional to the rate of piston movement.

[0118] In one or more embodiments, the actuator 1305 may be a linear actuator. Alternatively, the actuator may be a rotary actuator, as shown in the exemplary system 1402 of FIG. 14. The system 1402 is otherwise similar to the system 1302 of FIG. 13 in that it includes a damper 1425 connected between the aircraft boom 1405 and a prop rotor frame 1422, which is tiltably connected to a fixed frame 1420 such that the prop rotor frame 1422 (and the prop rotor) can tilt about the axis of rotation 1401. The system 1402 includes a rotary actuator 1406 connected to the prop rotor frame 1422 that drives the tilting of the prop rotor about the axis of rotation 1401. In the illustrated embodiment, the rotary actuator 1406 includes an electric motor connected to a gear train that drives a worm gear. The worm gear drives the gear to rotate about the axis of rotation 1401, thereby tilting the prop rotor.

[0119] If actuator 1306 (or actuator 1406 in FIG. 14 ) becomes disconnected from prop rotor frame 1322 or 1422, respectively, a catastrophic failure may occur, as the prop rotor may begin to tilt rapidly without control. In one or more embodiments, system 1302 or 1402 may include a second (redundant) actuator also connected to prop rotor frame 1322 or 1422, such that even if first actuator 1306 or 1406 becomes disconnected, the second actuator can still control the tilting of the prop rotor about axis of rotation 1301 or 1401. However, adding a second actuator may complicate system 1302 or 1402, increase costs, and add more weight to the aircraft. Rather than implementing a second actuator, systems 1302 and 1402 instead include a damper mechanism, such as damper 1325 or 1425, respectively. If the actuator 1306 or 1406 becomes disconnected from the prop rotor frame 1322 or 1422, the damper 1325 or 1425 dissipates energy, thereby damping the tilting speed of the prop rotor, which eliminates the catastrophic consequences of the actuator 1306 or 1406 becoming disconnected from the prop rotor frame 1322 or 1422.

[0120] As described above, the damper 1325 or 1425 can be rotatably connected to a pin 1308 or 1408 that extends between two sides of the clevis bracket 1309 or 1409, which is fixedly attached to the prop rotor frame 1322 or 1422, respectively. If the actuator 1306 (or actuator 1406 in FIG. 14 ) is disconnected from the prop rotor frame 1322 or 1422, respectively, the damper 1325 or 1425 will damp the tilting speed of the prop rotor when the prop rotor is moving but will not apply a damping force when the prop rotor is stationary. As the prop rotor tilts toward a forward flight position, the prop rotor frame 1322 or 1422 exerts a compressive force on the damper piston rod 1307 or 1407 via the clevis bracket 1309 or 1409, urging the damper piston rod 1307 or 1407 to retract into the cylinder 1316 or 1416. Similarly, the prop rotor frame 1322 or 1422 exerts a tensile force on the damper piston rod 1307 or 1407 via the clevis bracket 1309 or 1409, urging the piston rod 1307 or 1407 to extend outward from the piston cylinder 1316 or 1416 as the prop rotor frame 1322 or 1422 (and the prop rotor) tilts toward a raised position. As the prop rotor frame 1322 or 1422 (and the prop rotor) tilts in either direction after the actuator is disconnected, the fluid in the cylinder acts to resist extension or retraction of the damper piston rod 1307 or 1407 by exerting a reaction force acting on the piston 1318 or 1418 in the damper 1325 that is proportional to the rate of piston movement. Multiple openings in the piston allow damping fluid to flow through the piston so that the piston can move within the cylinder, thus allowing the damper piston rod 1307 or 1407 to extend or retract at a damped rate. The damper 1325 or 1425 thus dissipates the rotational energy of the prop rotor by exerting a force in a direction opposite the direction of tilt of the prop rotor.

[0121] As described above with respect to systems 1302 and 1402, dampers 1325 and / or 1425 may include balanced hydraulic or pneumatic cylinders 1316 and 1416, respectively, including pistons 1318 and 1418 sliding within the cylinders and connected to piston rods 1307 and 1407, respectively, as known in the art. The pistons include a plurality of openings (not shown) through which fluid flows as the pistons move within the cylinders. Balanced dampers contain less damping fluid than unbalanced dampers and therefore generally weigh less than unbalanced dampers. Balanced dampers disclosed herein may require 0.25 to 0.5 quarts of damping fluid. In some embodiments, balanced dampers require less than 0.35 quarts of damping fluid. In contrast, conventional dampers may require 0.51 to 1.0 quarts of damping fluid. In some embodiments, conventional dampers disclosed herein may require at least 0.74 quarts of damping fluid. In some embodiments, damper 1325 or 1425 can include a fire-resistant damping fluid, such as MIL-PRF-83282 or MIL-PRF-87257. MIL-PRF-83282 is a fire-resistant damping fluid, but the fluid becomes more viscous at low temperatures, causing more parasitic drag during cold operation. MIL-PRF-87257 maintains similar fire resistance to MIL-PRF-83282, but with improved low-temperature viscosity performance.

[0122] As described above, dampers 1325 and 1425 are passive dampers configured to apply a force (hydraulic or pneumatic) to propeller frame 1322 or 1422 (and propeller) respectively only when the propeller tilt angle is being adjusted (i.e., no force is applied when the propeller is stationary). In one or more embodiments, systems 1302 and 1402 may each include one or more redundant actuators and damper mechanisms such as damper 1325 or 1425.

[0123] The damper 1325 or 1425 can be configured to limit the rate of change of the tilt angle in both tilt directions of the prop rotor frame 1322 or 1422, respectively. The damper 1325 or 1425 can be positioned such that the force vector of the damper 1325 or 1425 can extend below the tilt axis, such as tilt axis 118 in FIG. 2 . In one or more embodiments, the damper 1325 or 1425 can be configured to limit the rate of change of the prop rotor tilt angle when the actuator 1306 or 1406, respectively, is disconnected from the prop rotor frame 1322 or 1422. According to some embodiments, the dampers described herein can limit the rate of change of the prop rotor tilt at maximum operating load to a maximum of 100 degrees per second. In some embodiments, the dampers can limit the rate of change of the prop rotor tilt angle to between 100 degrees per second and 300 degrees per second. In some embodiments, the damper may limit the rate of change of the prop rotor tilt angle to between 100 degrees per second and 200 degrees per second. In some embodiments, the damper may limit the rate of change of the prop rotor tilt angle to between 150 degrees per second and 180 degrees per second. In some embodiments, the damper may limit the rate of change of the prop rotor tilt angle to less than 200 degrees per second, less than 150 degrees per second, less than 100 degrees per second, less than 50 degrees per second, or less than 20 degrees per second.

[0124] As discussed above, in some embodiments, the damper 1325 or 1425 may be a balanced, unfilled damper. In such embodiments, the damper 1325 or 1425 may be configured to handle a peak damping pressure of 2000 psi to 4000 psi. In some embodiments, the damper 1325 or 1425 may be configured to handle a peak damping pressure of approximately 3000 psi. The damper 1325 or 1425 may be configured with an outer cylinder diameter of 1 to 2 inches, a piston diameter of 0.25 to 0.75 inches, a stroke of 9 to 11 inches, a body length of 10 to 12 inches, and an orifice diameter of 0.04 to 0.07 inches. The mass of the damper 1325 or 1425 may be 2 to 4 pounds. For example, an exemplary balanced unloaded damper may have a weight of approximately 2.2 pounds, which may include the damper cylinder, damper rod / piston, and damper fluid, as well as uncertainty factors for bearings, rod ends, trunnions, and other miscellaneous attachments. The peak power dissipated by the damper during transition (i.e., while the proprotor is transitioning between a forward flight configuration and a climb / hover configuration during normal operation) may be 0.7 watts, and the peak power dissipated by the damper during hover (i.e., when stabilizing and adjusting the proprotor tilt angle while the proprotor is in a climb / hover configuration during normal operation) may be 6.3 watts.

[0125] In other embodiments, damper 1325 or 1425 may be an unbalanced damper. In such embodiments, damper 1325 or 1425 may be configured to handle a maximum damping pressure of 2000 psi to 4000 psi, and in some embodiments, may be configured to handle a maximum damping pressure of 3000 psi. The unbalanced damper may be configured to operate at an actuation pressure (or precharge) of 1000 psi to 2000 psi, and in some embodiments, the actuation pressure / precharge may be configured at approximately 1500 psi. The load from the differential region of unbalanced damper 1325 may be 300 lbf to 400 lbf, and in some embodiments, the load from the differential region of unbalanced damper 1325 or 1425 may be approximately 330 lbf. The damper 1325 or 1425 may be configured with an outer cylinder diameter of 1.5 to 2.5 inches, an inner rod diameter of 0.25 to 0.75 inches, a stroke of 9 to 11 inches, a body length of 13 to 14 inches, and an orifice diameter of 0.08 to 0.10 inches. The mass of the damper 1325 or 1425 may be 2 to 4 pounds. For example, an exemplary unbalanced damper may weigh approximately 2.9 pounds, which may include the damper cylinder, damper rod / piston, and damper fluid, as well as uncertainty factors for bearings, rod ends, trunnions, and other miscellaneous attachments. The maximum power consumed by the damper 1325 or 1425 during transition may be 25 to 30 watts, and the peak power consumed by the damper 1325 during hover may be 60 to 65 watts.

[0126] According to the exemplary design specifications for the respective balanced and unbalanced dampers described above, the balanced and unloaded damper may offer advantageous configurations in terms of power consumption, primarily due to its lack of preload, and may have a smaller mass than the unloaded damper. Various additional or different design modifications may be made to the exemplary dampers described herein (e.g., dampers 1325 and 1425 described above and dampers 1525 and 1625 described below) without departing from the scope of the claims. For example, the passive damper may be configured for stroke-sensitive damping, which may be achieved by one or more axial grooves in the damper cylinder to reduce damping over a portion of the stroke, or by a stepped bore / piston with a secondary flow path with additional restriction. The passive damper may also be configured with a staged valve, an increased-rate shim stack, and / or pressure relief within the staged valve to allow for a wider range of speeds near the design operating load. Additionally, the passive damper may include a solenoid directional control valve with pressure relief. Passive dampers may also include various design features to mitigate potential failure of the damper. For example, the damper may include a sight glass for inspecting oil or a pressure transducer configured to monitor and verify nominal damping at normal operating speeds toward the aircraft endstops.

[0127] In one or more embodiments, controlling an aircraft including one or more dampers as discussed above includes receiving a command at a controller to adjust a tilt angle of a tiltable prop rotor tiltable between a climb position to provide lift to the aircraft and a forward flight position to provide forward thrust to the aircraft, and controlling at least one actuator to adjust the tilt angle of the tiltable prop rotor according to the command, wherein the at least one passive damper is coupled to the tiltable prop rotor to limit a rate of change of the tilt angle of the tiltable prop rotor.

[0128] Figures 15A and 15B illustrate a system 1502 mounted between an aircraft boom and a tiltable prop rotor, similar to systems 1302 and 1402 illustrated above in Figures 13 and 14. The tiltable prop rotor (omitted from Figures 13 and 14) is shown in a raised position in Figure 15A and a forward flight position in Figure 15B. Figures 15A and 15B therefore depict how the damper 1525 moves in response to changes in the tilt angle of the tiltable prop rotor.

[0129] FIG. 15A depicts a first configuration of the system 1502, including a damper 1525. The system 1502, omitted from FIG. 13 , includes a prop rotor 1503 in a raised configuration, rotatably connected to both a linear damper 1525 and a linear actuator 1506. The damper 1525 and actuator 1506 of the system 1502 are connected between a boom 1505 of an aircraft, such as the aircraft 100 of FIGS. 1 and 2 , and a prop rotor frame 1522 to which the prop rotor 1503 is mounted. In one or more embodiments, the boom 1505 can include a rib 1550, and the actuator 1506 and damper 1525 can be positioned on either side of the rib 1550. The boom 1505 can also include one or more bulkheads 1555 through which a portion of the piston rod 1507 of the damper 1525 extends. The prop rotor frame 1522 is tiltably connected to the fixed frame 1520 such that the prop rotor frame 1522 (and the prop rotor 1503) can tilt about the axis of rotation 1501. The actuator 1506 can connect to the prop rotor frame 1522 and drive the tilting of the prop rotor about the axis of rotation 1501. In one or more embodiments, the actuator 1505 can be a linear actuator.

[0130] The actuator 1506 can be configured to tilt the prop rotor 1503 between a climb configuration depicted in FIG. 15A and a forward flight configuration depicted in FIG. 15B. When the actuator tilts the prop rotor 1503 between the climb configuration depicted in FIG. 15A and the forward flight configuration depicted in FIG. 15B, the damper 1525 can exert a force in an opposing direction to the movement of the prop rotor. As the actuator tilts the prop rotor 1503, fluid within the cylinder 1516 of the damper 1525 acts to resist extension or retraction of the damper piston rod 1507 by exerting an opposing force acting on a piston (not shown) within the cylinder 1516 of the damper 1525 that is proportional to the rate of piston movement. Multiple openings (not shown) in the piston allow damping fluid to flow through the piston so that the piston can move within the cylinder. 13 and 14, the damper 1525 is a passive damper configured to apply a force (hydraulic or pneumatic) to the prop rotor frame 1522 (and the prop rotor) only when the tilt angle of the prop rotor is changing (i.e., no bias is applied when the prop rotor is stationary).

[0131] As described above, without the damper 1525, if the actuator 1506 were to become disconnected from the prop rotor frame 1522 (e.g., by an actuator shaft break), the prop rotor 1503 would rapidly tilt in an uncontrolled manner, potentially leading to catastrophic failure of the aircraft. The damper 1525 is configured to control the rate of change of the tilt angle of the prop rotor 1503 when the actuator 1506 becomes disconnected from the prop rotor frame to prevent the prop rotor 1503 from tilting at a rate that could damage the aircraft. The damper 1525 is configured to apply a compressive or tensile force to the prop rotor frame 1522, depending on the direction of tilt, in substantially the same manner as when the actuator is connected to the prop rotor, thus damping the tilting rate of the prop rotor when the actuator becomes disconnected.

[0132] Figure 16 illustrates a system 1602 including a damper in an alternative configuration to those shown in Figures 13, 14, and 15A-15B. In one or more embodiments, any of the systems described above, such as system 202, system 302, system 402, system 502, and aircraft 100, can include a damper to limit the tilt rate of the prop rotor, such as in the event of an actuator failure, as shown in example system 1602 of Figure 16. System 1602 can be implemented in aircraft having tiltable prop rotors in place of, or in addition to, redundant actuators and / or joint assemblies including failsafe latches, as described below.

[0133] The damper 1625 of the system 1602 is connected between a stationary frame 1620 of a boom 1605 of an aircraft, such as the aircraft 100 of FIGS. 1 and 2 , and a prop rotor frame 1622 to which a prop rotor (not shown) is mounted. The actuator 1606 may also be connected between the stationary frame 1620 and the prop rotor frame 1622 to drive tilting of the prop rotor about the axis of rotation 1601. In one or more embodiments, the actuator 1605 may be a linear actuator. The damper 1625 and the actuator 1606 may both be rotatably connected to the stationary frame 1620 of the boom 1605 and to the prop rotor frame 1622 such that the prop rotor frame (and prop rotor) can rotate relative to the damper 1625 and the actuator 1606.

[0134] Similar to dampers 1325 and 1425, if actuator 1606 becomes disconnected from prop rotor frame 1622 (e.g., due to actuator shaft breakage), catastrophic failure can occur as the prop rotor can begin to tilt rapidly without control. In one or more embodiments, system 1602 can include a second (redundant) actuator also connected to prop rotor frame 1622 so that even if first actuator 1606 becomes disconnected, the second actuator can still control tilting of the prop rotor about axis of rotation 1601. However, as described throughout, adding a second actuator can complicate system 1602, increase costs, and add more weight to the aircraft. Rather than implementing a second actuator, system 1602 instead includes a damper mechanism, such as damper 1625. If the actuator 1606 becomes disconnected from the prop rotor frame 1622, the damper 1625 dissipates energy, thereby limiting the rate of change of the prop rotor tilt (e.g., as described above via a damping fluid exerting a force on the damper piston), which eliminates the catastrophic consequences of the actuator 1606 becoming disconnected from the prop rotor frame 1622.

[0135] Similar to dampers 1325, 1425, and 1525, damper 1625 may be rotatably connected to a pin (not shown) extending between two sides of a clevis bracket 1609 fixedly attached to the prop rotor frame 1622. When the actuator 1606 is disconnected from the prop rotor frame 1622, the damper 1625 exerts a compressive force on the prop rotor frame 1622 through the clevis bracket 1609 as the prop rotor frame 1622 (and the prop rotor) tilts toward a forward flight position, and exerts a tensile force on the prop rotor frame 1622 through the clevis bracket 1609 as the prop rotor frame 1622 (and the prop rotor) tilts toward a raised position. Thus, the damper 1625 dissipates energy by exerting a force in a direction opposite to the direction of tilt of the prop rotor.

[0136] The damper 1625 may be configured to handle a maximum operating pressure of 2000 psi to 4000 psi. In some embodiments, the maximum operating pressure of the damper 1625 may be approximately 3000 psi. The outer diameter of the cylinder 1616 of the damper 1625 may be 1.0 to 2.5 inches. In some embodiments, the outer diameter of the damper 1625 may be approximately 2.0 inches. The piston diameter 1607 of the damper 1625 may be 0.25 to 0.75 inches. In some embodiments, the diameter of the piston 1607 may be approximately 0.5 inches. The normal operating stroke of the damper 1625 may be 3 to 6 inches. In some embodiments, the normal operating stroke of the damper 1625 may be approximately 4.65 inches. The pin-to-pin range of the normal operating stroke of the damper 1625 may be 9 to 16 inches. In some embodiments, the pin-to-pin range of the normal operating stroke of the damper 1625 may be 9.9 to 14.6 inches. The mechanical stroke of the damper 1625 can be 3 to 6 inches. In some embodiments, the mechanical stroke of the damper 1625 can be a minimum of approximately 5.06 inches. The pin-to-pin range of the mechanical stroke can be 9 to 15 inches. In some embodiments, the pin-to-pin range of the mechanical stroke can be 7.7 to 14.8 inches.

[0137] In one or more embodiments, the damper (e.g., damper 1325, 1425, 1525, or 1625) may be configured to generate a worst-case scenario hinge moment less than the maximum allowable impact velocity at an end stop of the aircraft, such as end stop 1548 shown in FIG. 15 , which is a component configured to be impacted by a tiltable component, such as a prop rotor, at the maximum tilt degree in each direction of the aircraft, as described further below. In some embodiments, the worst-case hinge moment at the maximum allowable impact velocity may be 2500 lbf to 3000 lbf at a damper piston linear velocity of about 12 inches per second. In this manner, the damper may be configured to provide a varying amount of force depending on the speed of linear extension or retraction of the actuator to tilt the tiltable prop rotor during normal operation.

[0138] An exemplary graph showing the damping force of a damper using TPD MD-914-04 damping fluid at various rates of linear extension and retraction is shown in FIG. 17. The linear velocity of the exemplary damper, located on the X-axis of the graph in FIG. 17, is the rate of extension or retraction of one of the exemplary damper pistons, e.g., dampers 1325, 1425, 1525, and 1625. Positive values ​​on the X-axis indicate extension of the damper piston (e.g., damper piston 1318 or 1418), and negative values ​​on the X-axis indicate retraction of the damper piston. As an actuator (e.g., actuator 1306 or 1406) configured to tilt a prop rotor (e.g., prop rotor 203 in FIG. 2) tilts the prop rotor toward a raised configuration (or, in the event of actuator shaft failure, the prop rotor tilts toward the raised configuration in an uncontrolled manner), the damper exerts a force acting in tension on the prop rotor proportional to the piston velocity. Alternatively, as the actuator tilts the prop rotor toward a forward flight configuration (or, again, if the actuator shaft breaks, the prop rotor tilts in an uncontrolled manner toward a climb configuration), the damper exerts a compressive force on the prop rotor. Thus, the damper exerts a force in a direction opposite the direction of prop rotor tilt while the prop rotor is tilting.

[0139] FIG. 17 includes three different lines 1701-1703 illustrating three different relationships between the linear velocity of damper piston extension or retraction and damping force. Line 1701 illustrates the lower threshold damping force at various linear velocities, line 1702 illustrates the nominal damping force at various linear velocities for an exemplary damper, and line 1703 illustrates the upper threshold damping force at various linear velocities. As shown in FIG. 17, the exemplary damper can be configured to maintain a damper force of less than 400 lbf at actuator extension and retraction speeds of + / - 0.68 inches per second to minimize losses to the actuator. The exemplary damper using MIL-PRF-87257 damping fluid depicted in FIG. 17 is configured to handle a limit load of 5000 lbf in extension and 3000 lbf in retraction.

[0140] The selection of a damper for damping high-speed tilting of a tiltable propeller when an actuator is disconnected can be based at least in part on the characteristics of the propeller. According to some embodiments, the tiltable propellers disclosed herein can have a mass moment of inertia of 5 to 10 kilograms per square meter. An exemplary propeller having a mass moment of inertia of approximately 6.7 kilograms per square meter was used to determine upper and lower limits on the rotational energy of the propeller upon impact with an aircraft end stop during uncontrolled tilting of the tiltable propeller, as described below.

[0141] According to some embodiments, the rotational energy of a tiltable prop rotor during uncontrolled rotation can be between 0 J and 1000 J upon impact with an aircraft endstop. A lower limit for rotational energy can be between 0 J and 300 J, and an upper limit for rotational energy can be between 700 J and 1000 J. The lower limit can include an instance where the actuator disengages and the tiltable prop rotor impacts the aircraft endstop at cruise speed after minimal acceleration (e.g., between 0 ms and 100 ms of acceleration). The upper limit can include an instance where the tiltable prop rotor accelerates after disengagement until the actuator disengages and impacts the aircraft endstop.

[0142] In some embodiments, uncontrolled tilting of a tiltable prop rotor due to actuator disconnection can result in the tiltable prop rotor impacting an aircraft end stop at 500 to 1000 degrees per second. According to some embodiments, the allowable impact velocity of the tiltable prop rotor impacting an aircraft end stop is 150 to 200 degrees per second, far below the 500 to 1000 degrees per second experienced during uncontrolled tilting. In some embodiments, the allowable impact velocity can be less than 100 degrees per second. Thus, the passive dampers provided herein can be configured to reduce the impact velocity of the tiltable prop rotor by preventing the prop rotor from accelerating beyond the allowable impact velocity in the event of actuator disconnection. The passive damper can be configured to reduce the impact velocity to less than 200 degrees per second, less than 150 degrees per second, less than 100 degrees per second, less than 50 degrees per second, or less than 20 degrees per second, according to various aircraft requirements. The rate at which the actuator needs to tilt the tiltable prop-rotor between the climb position and the forward flight position during normal operation can affect the amount by which the damper is configured to reduce the angular velocity of the prop-rotor. In some embodiments, the actuator is configured to tilt the tiltable prop-rotor between the climb position and the forward flight position at a maximum of 5 to 30 degrees per second. In some embodiments, the actuator is configured to tilt the tiltable prop-rotor between the climb position and the forward flight position at a maximum of 15 degrees per second. In some embodiments, the actuator is configured to tilt the tiltable prop-rotor at a maximum of 15 degrees per second during hover (i.e., when stabilizing adjustments are made to the prop-rotor tilt angle while the prop-rotor is in the climb / hover configuration during normal operation) and at a maximum of 8 degrees per second during transition (i.e., while the prop-rotor is transitioning between the climb / hover configuration and the forward flight configuration during normal operation).

[0143] The results of an exemplary tiltable prop rotor shear simulation analysis are shown in Figures 18A-18D. It should be understood that the simulation results depicted in Figures 18A-18D are only meant to provide exemplary performance data for exemplary passive damping systems. The results of the tiltable prop rotor shear simulation analysis should not be construed as limiting, as one skilled in the art would understand that many different passive dampers could be configured to damp the tilting speed of a tiltable prop rotor described herein without departing from the scope of the claims.

[0144] FIG. 18A depicts both the external hinge moment and the damped hinge moment acting on the prop rotor as a function of time beginning approximately at the instant of actuator disconnection. The hinge moment, provided on the y-axis of the graph in FIG. 18A, is the moment acting about the prop rotor's axis of rotation about the joint connecting the prop rotor to the aircraft 100. The prop rotor's axis of rotation is shown in the example of FIG. 13 as axis of rotation 1301. FIG. 18A includes two lines, lines 1801 and 1802. Line 1801 represents the prop rotor hinge moment, and line 1802 shows the damper hinge moment. As illustrated, the damper hinge moment can closely track the prop rotor hinge moment. Within 200 milliseconds of actuator disconnection, the hinge moments of both the prop rotor and the damper can increase from 400 ft-lbf to 600 ft-lbf. The hinge moment may then temporarily decrease before rising to a maximum value of 800 ft-lbf to 1000 ft-lbf 1 to 2 seconds after the actuator is disconnected, as shown.

[0145] FIG. 18B depicts a plot of prop rotor angular velocity as a function of time beginning approximately at the instant of actuator disconnection. The angular velocity provided on the Y-axis of the graph in FIG. 18B is the angular velocity of the prop rotor rotating about its axis of rotation, which is shown in the example of FIG. 13 as axis of rotation 1301. The angular velocity depicted on the Y-axis is the angular velocity of the prop rotor after actuator disconnection acting under the influence of a damping force provided by a passive damper, such as, for example, damper 1325, 1425, 1525, or 1625. As shown, the angular velocity of the prop rotor after actuator disconnection can be configured to remain below an impact velocity of 100 degrees per second when the angular velocity is controlled by a damper, for example, damper 1325, 1425, 1525, or 1625.

[0146] FIG. 18C depicts a plot of the linear velocity of the damper piston as a function of time beginning approximately at the instant of actuator disconnection. The linear velocity provided on the Y-axis of the graph in FIG. 18C is the linear velocity of an exemplary damper piston (e.g., of damper 1325 or 1425) traveling along the axis of damper extension and retraction. The approximate axis of damper extension and retraction is shown as axes 1311 and 1411 in FIGS. 13 and 14. As shown, immediately after the actuator disconnects (within 200 milliseconds), the damper piston can reach a linear velocity of 2 inches per second to 3 inches per second. Within the first 2 seconds after actuator disconnection, the linear velocity of the damper piston can vary from 2.0 to 3.5 inches per second.

[0147] FIG. 18D depicts a plot of damper force as a function of time beginning approximately at the instant of actuator disconnection. The damper force, provided on the Y-axis of the graph in FIG. 18D, is the force acting against the rotation of the prop rotor after actuator disconnection. As shown, immediately after the actuator disconnects (within 200 milliseconds), the damper force can reach an initial peak of 2500 lbf to 3500 lbf. Within the first two seconds after actuator disconnection, the damper force can fluctuate between 1500 lbf and 5000 lbf.

[0148] 19 and 20 illustrate plots of exemplary damper effective moment arm lengths as a function of propeller rotor tilt angle. The exemplary damper in FIG. 19 is mounted in a first exemplary configuration, and the exemplary damper in FIG. 20 is mounted in a second exemplary configuration. The effective moment arm lengths of the exemplary damper mounted in the first configuration, provided on the y-axis of the graph in FIG. 19, are the moment arm of the damper at various tilt angles of the tiltable propeller rotor from 0 degrees to 100 degrees. The effective moment arm lengths of the exemplary damper mounted in the second exemplary configuration, provided on the y-axis of the graph in FIG. 20, are the moment arm of the damper at various tilt angles of the tiltable propeller rotor from 0 degrees to 100 degrees. As shown in FIGS. 19 and 20, the mounting configuration of the damper relative to the propeller rotor can affect the effective moment arm of the damper as the propeller rotor tilts to different tilt angles.

[0149] In one or more embodiments, any of the above-described systems, such as system 202, system 302, system 402, system 502, system 1302, system 1402, and aircraft 100, can include a tilt rotor locking mechanism, as shown in example system 2102 in Figure 21, which shows a side view of system 2102 in a first configuration, in accordance with one or more embodiments of the present disclosure. System 2102 can be implemented in an aircraft having a tiltable prop rotor and can lock the tilt of the prop rotor in place, instead of, or in addition to, redundant actuator and / or damper mechanisms.

[0150] The system 2102 is positioned between the aircraft boom 2105 and the prop rotor 2103 and can include a tension spring 2145 connected to a pulley 2140 via a connector 2160, which is also connected to a pawl 2152 via a connector 2142. The pawl 2152 can selectively engage with a sector gear 2150 based on movement of the pulley 2140, the pawl 2152 being configured to move toward the sector gear 2150 as the pulley 2140 moves away from the actuator 2106. The pulley 2140 is coupled to the actuator 2106 in normal operation.

[0151] If the actuator 2106 and pulley 2140 become disconnected (e.g., if the actuator 2106 becomes disconnected from the prop rotor 2103), the pulley 2140 will move away from the actuator 2106 due to the bias from the spring 2145, as shown in the configuration of the system 2102 shown in FIG. 22. While FIG. 21 shows the pawl 2152 disengaged from the sector gear 2150 and the pulley 2140 located adjacent to the actuator 2106, FIG. 22 shows the pawl 2152 engaged with the sector gear 2150 and the pulley 2140, with the pulley 2140 no longer located adjacent to the actuator 2106. The tension spring 2145 can automatically pull the pulley 2140 away from the actuator 2106 if the actuator 2106 and pulley 2140 become disconnected, thereby forcing the pawl 2152 into engagement with the sector gear 2150. When the pawl 2152 engages the sector gear 2150, the prop rotor can be prevented from tilting further in one or both directions. By preventing further tilting in one or both directions, the system 2102 can prevent catastrophic failure if the actuator disconnects from the prop rotor without requiring redundant actuators or damping mechanisms. Figure 23 shows a front view of the example system of Figure 21 in accordance with one or more embodiments of the present disclosure.

[0152] In one or more embodiments, the sector gear 2250 can be a ratchet having a ridge that contacts a pawl 2252. Such an exemplary configuration is shown in FIG. 24A , which illustrates an exemplary ratchet configuration 2202, according to one or more embodiments of the present disclosure. The ratchet configuration 2202 includes a sector gear 2250 having multiple ridges 2253 with angled and straight sides. The pawl 2252 overcomes the angled sides but hooks onto the straight sides. In this manner, the prop rotor can tilt in a first direction (e.g., allowed to tilt counterclockwise in the illustrated configuration) but cannot tilt in a second direction (e.g., not allowed to tilt clockwise in the illustrated configuration). This can be useful for allowing the prop rotor to move to a desired fault-state tilt angle, such as a climb configuration, where the prop rotor can still be used during at least a portion of flight. Such a ratchet mechanism can also be used with any of the damper configurations described above to provide a slower tilt rate in the ratchet direction.

[0153] Alternatively, the sector gear and pawl can be configured to lock the prop rotor in both directions. Figure 24B illustrates an example locking configuration 2204 having a sector gear 2250 with multiple ridges 2255 with two straight sides and a locking mechanism 2254 that engages with the ridges 2255. When the locking mechanism 2254 is pressed against the gear sector 2250, the locking mechanism engages with the ridges 2255. Due to their straight sides, the locking mechanism 2254 cannot ride past the ridges 2255 in either direction, locking the sector gear 2250 (and thereby the prop rotor) in place.

[0154] Accordingly, described herein are systems and methods for mechanically linking the tilt of an aircraft propeller with the pitch of the propeller's blades. The system enables blade pitch to be adjusted for different operating regimes of the propeller while avoiding the need for, and the cost, weight, and failure points associated with, dedicated blade pitch adjustment actuators.

[0155] FIG. 25 illustrates an alternative spring configuration and function to provide a "mid-life" check function to ensure the ratchet mechanism does not jam.

[0156] For purposes of explanation, the foregoing description has been described with reference to specific embodiments. However, the foregoing illustrative discussion is not intended to be exhaustive or to limit the invention to the precise form disclosed. Many modifications and variations are possible in light of the above teachings. The embodiments were chosen and described in order to best explain the principles of the technology and their practical application, thereby enabling those skilled in the art to best utilize the technology and various embodiments with various modifications as suited to the particular use contemplated.

[0157] Although the disclosure and examples have been fully described with reference to the accompanying drawings, it should be noted that various changes and modifications will become apparent to those skilled in the art, and such changes and modifications are to be understood as being included within the scope of the disclosure and examples as defined by the claims.

Claims

1. A tilt device for an aircraft, comprising: a tiltable aircraft component that tilts between a first position and a second position; at least one actuator for adjusting the tilt angle of the tiltable aircraft component; at least one passive damper connected to the tiltable aircraft component and configured to limit a rate of change of the tilt angle of the tiltable aircraft component; a tilting apparatus, wherein the at least one passive damper is connected to the tiltable aircraft component on a first load path, and the at least one actuator is connected to the tiltable aircraft component on a second load path separate from the first load path of the at least one passive damper.

2. A tilting device as described in claim 1, wherein the tiltable aircraft component comprises a prop rotor.

3. A tilting device as described in claim 1, wherein the tiltable aircraft component comprises a prop rotor frame.

4. A tilting device as described in claim 1, wherein the at least one passive damper is configured to limit the rate of change of the tilt angle of the tiltable aircraft component in both tilt directions.

5. The tilting device according to any one of claims 1 to 4, wherein the at least one passive damper includes a hydraulic cylinder or a pneumatic cylinder.

6. 6. The tilting device of claim 5, wherein the hydraulic or pneumatic cylinder is a balanced unfilled cylinder or an unbalanced cylinder.

7. The tilting device of any one of claims 1 to 4, wherein the at least one actuator comprises a linear actuator or a rotary actuator.

8. A tilting device described in any one of claims 1 to 4, wherein the tiltable aircraft component is tiltably mounted on a boom.

9. A tilt device as described in any one of claims 1 to 4, wherein the at least one passive damper is at least partially contained within the outer shell.

10. A tilt device as described in any one of claims 1 to 4, wherein the at least one actuator and the at least one passive damper are positioned on both sides of a rib of the aircraft.

11. 5. The tilting apparatus of claim 1, wherein the tiltable aircraft component is configured to tilt in an upward direction from a forward flight position to a raised position, and a force vector of the at least one passive damper extends below a tilt axis of the tiltable aircraft component.

12. The tilting device according to any one of claims 1 to 4, wherein the tilt angle range of the tiltable aircraft component is at least 90 degrees.

13. A tilting device as described in any one of claims 1 to 4, further comprising a single actuator for adjusting the tilt angle of the tiltable aircraft component.

14. 5. The tilting apparatus of claim 1, wherein the at least one passive damper is configured to limit the rate of change of the tilt angle to a predetermined threshold when the tiltable aircraft component is disconnected from the at least one actuator during forward flight.

15. A tilt device as described in any one of claims 1 to 4, wherein the first position is an ascending position for providing an ascending force to the aircraft, and the second position is a forward thrust position for providing a forward thrust force to the aircraft.

16. The tilting device according to any one of claims 1 to 4, wherein the tiltable aircraft component is electrically driven.

17. An aircraft, Wings and a boom attached to the wing; at least one tiltable aircraft component mounted on the boom that tilts between a first position and a second position; at least one actuator for adjusting the tilt angle of the at least one tiltable aircraft component; at least one passive damper connected to the at least one tiltable aircraft component and configured to limit a rate of change of the tilt angle of the at least one tiltable aircraft component; the at least one passive damper is on a load path separate from a load path of the at least one actuator.

18. An aircraft as described in claim 17, wherein the at least one tiltable aircraft component is mounted to the forward end of the boom.

19. 19. An aircraft according to claim 17 or 18, wherein the booms are mounted inboard of the wing tips.

20. The at least one tiltable component comprises a plurality of tiltable components; 19. An aircraft as claimed in claim 17 or 18, wherein the at least one passive damper comprises a plurality of dampers.

21. 19. An aircraft according to claim 17 or 18, wherein the aircraft is a passenger aircraft.

22. 1. A method for controlling a tiltable component of an aircraft, comprising: receiving a command at a controller to adjust a tilt angle of a tiltable aircraft component between a first position and a second position; controlling at least one actuator to adjust the tilt angle of the tiltable aircraft component; at least one passive damper connected to the tiltable aircraft component and configured to limit a rate of change of the tilt angle of the tiltable aircraft component; the at least one passive damper is connected to the tiltable aircraft component on a first load path, and the at least one actuator is connected to the tiltable aircraft component on a second load path separate from the first load path of the at least one passive damper.

23. The method of claim 22, wherein the tiltable aircraft component comprises a prop rotor.

24. The method of claim 22, wherein the at least one passive damper is configured to limit the rate of change of the tilt angle of the tiltable aircraft component in both tilt directions.

25. A method according to any one of claims 22 to 24, wherein the at least one passive damper comprises a hydraulic cylinder or a pneumatic cylinder.

26. A method according to any one of claims 22 to 24, wherein the at least one actuator and the at least one passive damper are positioned on either side of a rib of the aircraft.

27. 25. The method of any one of claims 22 to 24, wherein the tiltable aircraft component is configured to tilt upwardly from a forward flight position to a raised position, and a force vector of the at least one passive damper extends below a tilt axis of the tiltable aircraft component.

28. A method according to any one of claims 22 to 24, wherein the tilt angle range of the tiltable aircraft component is at least 90 degrees.

29. A method described in any one of claims 22 to 24, wherein the at least one passive damper is configured to limit the rate of change of the tilt angle to a predetermined threshold when the tiltable aircraft component is disconnected from the at least one actuator during forward flight.

30. A method described in any one of claims 22 to 24, wherein the first position is an ascending position for providing an ascending force to the aircraft, and the second position is a forward thrust position for providing a forward thrust force to the aircraft.

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