Redundant load path for electric lift tilt actuator

Abstract

The invention provides a system for an aircraft that includes a tiltable propeller tiltable between a lift position for providing lift for the aircraft and a forward flight position for providing forward propulsion for the aircraft. The system comprises: at least one actuator for adjusting the tilt angle of the tiltable paddle; and at least one passive damper connected to the tiltable paddle and configured to limit a rate of change of the tilt angle of the tiltable paddle.

Description

[0001] Cross - Reference to Related Applications

[0002] This application claims the benefit of U.S. Provisional Application No. 63 / 378,536, filed Oct. 6, 2022, U.S. Provisional Application No. 63 / 378,680, filed Oct. 7, 2022, and U.S. Provisional Application No. 18 / 296,062, filed Apr. 5, 2023, the entire contents of each of which are incorporated herein by reference. Technical Field

[0003] The present disclosure generally relates to aircraft, and more particularly, to aircraft having tiltable rotors. Background Art

[0004] A vertical takeoff and landing (VTOL) aircraft is an aircraft that can take off and land vertically and hover. To take off and land vertically and hover, a VTOL aircraft may include one or more rotors that can be tilted between a position for providing vertical thrust for takeoff and landing and hovering and a position for providing forward thrust for forward flight. A VTOL aircraft may include wings, as in a conventional fixed-wing aircraft, that provide lift during forward flight, where the tiltable rotors may be mounted to the wings. Because tiltable rotors control the performance of the aircraft in flight, they must be designed to avoid catastrophic failures. This generally may include introducing redundant actuators for controlling the tilt of the rotors such that if one actuator fails, another actuator takes over, and / or overdesigning the aircraft structural components such that a particular failure mode is not catastrophic. However, redundant actuators and overdesigned components introduce additional unwanted weight, which may increase the manufacturing cost and the operating cost of the aircraft. Summary of the Invention

[0005] Disclosed herein is a secondary load path including a passive damper for an actuated tiltable rotor of an aircraft, the secondary load path being configured to inhibit high-speed tilting of the tiltable rotor in the event that an actuator that tilts the tiltable rotor becomes disconnected. The shaft of the actuator for tilting the tiltable rotor may be a single point of failure that may experience relatively high loads. A break in the shaft may cause the rotor to tilt uncontrollably, which may result in relatively high rotational speeds and relatively high support structure impact speeds under load, resulting in a catastrophic failure. The passive damper described herein is configured to inhibit the rotational speed of the tiltable component in the event of actuator disconnection, thereby preventing catastrophic failure without the need for redundancy and overdesign of the support structure.

[0006] In some examples, a damper configured to inhibit high-speed tilting of a tiltrotor as described herein may include a balanced hydraulic or pneumatic cylinder ("balanced damper without resistive damping"). The balanced damper without resistive damping cylinder may include a piston that slides within the damper's cylinder, the piston including a plurality of holes through which damping fluid flows as the piston moves through the cylinder. A piston rod connected to the damper's piston may be rotatably connected to the tiltrotor such that when the rotor tilts toward a lift configuration or a forward flight configuration, the damper applies a force in a direction opposite to the rotation of the tiltrotor. The force is a result of the drag force that the damping fluid applies to the damper piston as the damper piston moves through the damping fluid. Thus, as the piston moves through the damping fluid, the damper dissipates energy (e.g., by converting kinetic energy into heat) due to the drag force applied to the piston by the damping oil. In some examples, the damper may include an unbalanced cylinder ("unbalanced damper"). The unbalanced damper similarly includes a piston that slides within the cylinder and includes a plurality of holes through which damping fluid flows as the piston moves through the cylinder, and a pressurized gas may also be included within the cylinder in addition to the damping fluid. When a force is applied to the damper piston, the pressurized gas may compress or expand, depending on the direction of the force. Thus, the pressurized gas may absorb an initial impact force applied to the damper piston.

[0007] According to one aspect, a tiltrotor of an aircraft that is tiltable between a lift position and a forward flight position is connected to a secondary load path and includes a passive damper such as the balanced damper without resistive damping or the unbalanced damper described above. In the lift position, the tiltrotor is configured to provide lift for the aircraft, for example, during vertical takeoff. In the forward flight position, the tiltrotor is configured to provide forward propulsion for the aircraft. The aircraft may include an actuator connected to the tiltrotor for adjusting the tilt angle of the tiltrotor between the lift position and the forward flight position. As described above, due to the high loads experienced by the actuator shaft, the shaft may break, resulting in a freely floating tiltrotor, which may lead to a catastrophic failure of the aircraft. Accordingly, the aircraft disclosed herein is provided with at least one passive damper that is connected to the tiltrotor and is configured to limit the rate of change of the tilt angle of the tiltrotor.

[0008] An exemplary aircraft is provided that includes: a tiltrotor that is tiltable between a lift position for providing lift for the aircraft and a forward flight position for providing forward propulsion for the aircraft; at least one actuator for adjusting the tilt angle of the tiltrotor; and at least one passive damper that is connected to the tiltrotor and is configured to limit the rate of change of the tilt angle of the tiltrotor.

[0009] In some examples of the aircraft, the at least one passive damper includes a hydraulic cylinder or a pneumatic cylinder.

[0010] In some examples of the aircraft, the hydraulic cylinder or pneumatic cylinder is a balanced non-powered cylinder or an unbalanced cylinder.

[0011] In some examples of the aircraft, at least one passive damper is configured to limit the rate of change of the tilt angle of the tiltable rotor in two tilt directions.

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

[0013] In some examples of the aircraft, at least one actuator includes a rotary actuator.

[0014] In some examples of the aircraft, the aircraft includes a boom, and the tiltable rotor is tiltably mounted to the boom.

[0015] In some examples of the aircraft, the aircraft includes a fixed rotor, and the fixed rotor is mounted to the boom at a fixed position to provide lift.

[0016] In some examples of the aircraft, the tiltable rotor is mounted to the front end of the boom, and the fixed rotor is mounted to the rear end of the boom.

[0017] In some examples of the aircraft, the boom includes a housing, and at least one passive damper is at least partially received within the housing.

[0018] In some examples of the aircraft, the boom includes ribs, and at least one actuator and at least one passive damper are positioned on opposite sides of the ribs.

[0019] In some examples of the aircraft, the boom is mounted to the wing of the aircraft, inside the end of the wing.

[0020] In some examples of the aircraft, the tiltable rotor is configured to tilt from a forward flight position to a lift position in an upward direction, and the force vector of at least one passive damper extends below the tilt axis of the tiltable rotor.

[0021] In some examples of the aircraft, the aircraft includes a plurality of tiltable rotors and a plurality of passive dampers for the plurality of tiltable rotors.

[0022] In some examples of the aircraft, the range of the tilt angle of the tiltable rotor is at least 90 degrees.

[0023] In some examples of the aircraft, the aircraft includes a single actuator for adjusting the tilt angle of the tiltable rotor.

[0024] In some examples of the aircraft, at least one passive damper is configured to limit the rate of change of the tilt angle to a predetermined threshold in the case where the tiltable rotor is disconnected from at least one actuator during forward flight.

[0025] In some examples of the aircraft, the tiltable propellers are electric.

[0026] In some examples of the aircraft, the aircraft is a passenger aircraft.

[0027] A method for controlling an aircraft is provided, the method comprising: receiving, at a controller, a command to adjust a tilt angle of a tiltable propeller, the tiltable propeller being tiltable between a lift position for providing lift to the aircraft and a forward flight position for providing forward propulsion to the aircraft; and controlling at least one actuator in accordance with the command to adjust the tilt angle of the tiltable propeller, wherein at least one passive damper is connected to the tiltable propeller to limit a rate of change of the tilt angle of the tiltable propeller. The method may be performed using any one of the above aircraft.

[0028] It should be understood that any variations, aspects, features, and options described in view of the system equally apply to the method, and vice versa. It will also be clear that any one or more of the above variations, aspects, features, and options may be combined. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] The present invention will now be described, by way of example only, with reference to the accompanying drawings, in which:

[0030] Figure 1 An aircraft in a forward flight configuration according to one or more examples of the present disclosure is shown;

[0031] Figure 2 For one or more examples of the present disclosure Figure 1 A perspective view of an aircraft, showing the propeller positions in the lift configuration and the forward flight configuration;

[0032] Figure 3 A side view of a portion of an aircraft having an exemplary system according to one or more examples of the present disclosure is shown;

[0033] Figure 4 A detailed view of an exemplary system according to one or more examples of the present disclosure is shown;

[0034] Figure 5 An example of one or more examples of the present disclosure is shown Figure 4 A detailed view of a portion of an exemplary system;

[0035] Figure 6 An example of one or more examples of the present disclosure is shown Figure 4 A detailed view of a portion of an exemplary system;

[0036] Figure 7Show a top perspective view and a detailed view of an exemplary system in accordance with one or more examples of the present disclosure;

[0037] Figure 8 Show a detailed view of an exemplary system with a tension control rod in accordance with one or more examples of the present disclosure;

[0038] Figure 9 Show an exemplary Figure 8 detailed cross-sectional view of a bushing and a cam interface of an exemplary system in accordance with one or more examples of the present disclosure;

[0039] Figure 10 Show a detailed cross-sectional view of an exemplary hub of a controllable pitch propeller in accordance with one or more examples of the present disclosure;

[0040] Figure 11 Show a graph of an example of the blade pitch varying with the controllable pitch propeller tilt angle;

[0041] Figure 12 A figure showing the blade pitch angle;

[0042] Figure 13 Show a detailed view of an exemplary system in accordance with one or more examples of the present disclosure;

[0043] Figure 14 Show a detailed view of an exemplary system in accordance with one or more examples of the present disclosure;

[0044] Figure 15A Show a detailed view of an exemplary tiltable controllable pitch propeller damping system in a first configuration in accordance with one or more examples of the present disclosure;

[0045] Figure 15B Show a detailed view of an exemplary tiltable controllable pitch propeller damping system in a second configuration in accordance with one or more examples of the present disclosure;

[0046] Figure 16 Show a detailed view of an exemplary tiltable controllable pitch propeller damping system in accordance with one or more examples of the present disclosure;

[0047] Figure 17 Show a graph of an example of the damping force varying with the linear rate of the damper piston in accordance with one or more examples of the present disclosure;

[0048] Figure 18A Show a graph of the controllable pitch propeller and the damping hinge moment varying with time in accordance with one or more examples of the present disclosure;

[0049] Figure 18B Show a graph of the controllable pitch propeller angular rate varying with time in accordance with one or more examples of the present disclosure;

[0050] Figure 18CShow a graph of the linear rate of a damper piston varying with time according to one or more examples of the present disclosure;

[0051] Figure 18D Show a graph of the damping force varying with time according to one or more examples of the present disclosure;

[0052] Figure 19 Show a graph of the exemplary damper effective arm length varying with the propeller blade tilt angle for a first damper configuration according to one or more examples of the present disclosure;

[0053] Figure 20 Show a graph of the exemplary damper effective arm length varying with the propeller blade tilt angle for a second damper configuration according to one or more examples of the present disclosure;

[0054] Figure 21 Show a side view of an exemplary system in a first configuration according to one or more examples of the present disclosure;

[0055] Figure 22 Show the exemplary system of FIG. 15 in a second configuration according to one or more examples of the present disclosure;

[0056] Figure 23 Show a front view of the exemplary system of FIG. 15 according to one or more examples of the present disclosure;

[0057] Figure 24A Show an exemplary ratchet configuration and Figure 24B Show an exemplary locking configuration for locking a propeller in at least one direction; and

[0058] Figure 25 Show a front view of an exemplary system according to one or more examples of the present disclosure. Detailed Description

[0059] In the following description of various examples, reference is made to the accompanying drawings, in which specific examples that can be practiced are shown by way of illustration. The description is presented to enable one of ordinary skill in the art to make and use the invention, and the description is provided in the context of a patent application and its claims. Various modifications to the described examples will be apparent to those of skill in the art, and the general principles herein can be applied to other examples. Accordingly, the invention is not intended to be limited to the examples shown, but is to be accorded the widest scope consistent with the principles and features described herein.

[0060] This document describes systems, devices, and methods for using a passive damper connected between a tiltable rotor of an aircraft and a support structure to suppress high-speed tilting of the tiltable rotor of the aircraft. If the actuator disconnects (e.g., the connecting portion of the actuator disconnects from the remainder of the actuator due to a break in the actuator shaft), the movement of the tiltable rotor will be uncontrolled, posing a risk of catastrophic failure. A break in the actuator shaft can cause the tiltable rotor to tilt uncontrollably, which can result in a relatively high rotational speed and thus a relatively high impact speed of the support structure under load, leading to catastrophic failure of the aircraft.

[0061] To mitigate such risks, an aircraft can be designed with redundant components, such as redundant actuators, and / or overdesigned components that can withstand uncontrolled movement, both of which increase the weight and cost of the aircraft design and increase the operating cost. In accordance with the principles herein, a secondary load path incorporating a passive damper can eliminate the need for redundant actuators and / or overdesign of aircraft structural components to mitigate uncontrolled movement of the tiltable rotor by suppressing high-speed tilting of the tiltable rotor in the event that the actuator controlling the tilting of the tiltable rotor disconnects from the tiltable rotor, as further described below.

[0062] According to various examples, a tiltable rotor of an aircraft that can be tilted between a lift position and a forward flight position can be connected to a passive damper as described above. In the lift position, the tiltable rotor is configured to provide lift for the aircraft, for example, during takeoff. In the forward flight position, the tiltable rotor is configured to provide forward propulsion for the aircraft. To adjust the angle of the tiltable rotor (e.g., between the lift position and the forward flight position), the aircraft can include an actuator between each tiltable rotor and the structure of the aircraft, such as a boom. Due to the high loads experienced by the actuator shaft, a failure mode that must be considered in aircraft design is a break in the actuator shaft, which can cause the component (tiltable rotor) to float freely, which in turn can lead to catastrophic failure of the aircraft due to a high-speed impact of the tiltable rotor on the aircraft structural components. For this reason, the aircraft disclosed herein is provided with at least one passive damper that is connected to each tiltable rotor and is configured to limit the rate of change of the tilt angle of the tiltable rotor.

[0063] The passive dampers described herein may include a balance hydraulic cylinder or a pneumatic cylinder (“balance damper without resistive damping”). The balance damper without resistive damping cylinder may include a piston that slides within the cylinder and includes a plurality of holes through which damping fluid flows as the piston moves through the cylinder. The piston may be attached to a piston rod extending from the cylinder and may be rotatably attached to a tiltable propeller of an aircraft. When the tiltable propeller rotates toward the lift position or the forward flight position, it will apply a tensile force or a compressive force to the piston rod, respectively, thereby pushing the piston rod away from or toward the piston cylinder. In turn, when the propeller tilts in either direction, the piston will be forced to slide within the cylinder. Thus, as the piston moves through the damping fluid, the damper dissipates (e.g., by converting kinetic energy into heat) energy due to the resistance applied by the damping fluid to the piston.

[0064] 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 a plurality of holes through which damping fluid flows as the piston moves through the cylinder. In addition to the damping fluid, the cylinder may also include a pressurized gas. When a force is applied to the damper piston, the pressurized gas may compress or expand, depending on the direction of the force. Thus, the pressurized gas may absorb the initial impact force applied to the damper piston. Therefore, if the actuator shaft breaks, the passive damper (balance damper without resistive damping or unbalanced damper) inhibits the high-speed tilting of the tiltable propeller, thereby preventing a catastrophic failure.

[0065] Thus, the systems, devices, and methods described herein provide a more cost-effective aircraft design and reduce the weight of the entire aircraft while meeting safety requirements. Additionally, the passive damper does not require activation, engagement, or other controls to perform its function of controlling the rate of change of the tilt of the tiltable component. Thus, it provides a reliable and simple solution for the actuator disconnection failure mode described herein.

[0066] As used herein, unless the context clearly indicates otherwise, the singular forms “a / an” and “the” are also intended to include the plural forms. It should be understood that the term “and / or” as used herein refers to and encompasses any and all possible combinations of one or more of the associated listed items. It should also be understood that the terms “includes,” “including,” “comprises,” and “comprising,” when used herein, specify the presence of the stated features, integers, steps, operations, elements, components, and / or units, but do not preclude the presence or addition of one or more other features, integers, steps, operations, elements, components, units, and / or groups thereof.

[0067] As used herein, the term "pivoting propeller" refers to a variable pitch propeller, where the direction of the propeller's thrust can be changed by varying the pitch angle of the propeller. For example, the pitch angle can be changed from an angle that provides at least a certain degree of vertical thrust (e.g., for vertical takeoff and landing) to an angle that provides at least a certain degree of horizontal thrust (e.g., for forward flight). As used herein, a pivoting propeller lift configuration refers to any pivoting propeller orientation where the pivoting propeller thrust primarily provides lift for the aircraft, and a pivoting propeller forward flight configuration refers to any pivoting propeller orientation where the pivoting propeller thrust primarily provides forward thrust for the aircraft.

[0068] As used herein, "vertical takeoff and landing" ("VTOL") refers to the ability of an aircraft to move substantially vertically without relying solely on lift provided by the wings of the aircraft. While this term encompasses direct vertical takeoff and landing (i.e., vertical movement without any horizontal movement), it also encompasses 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 various examples described herein may take off and land in a conventional aircraft manner using lift provided by its wings at a suitable airspeed.

[0069] Figure 1 An aircraft 100 in a forward flight configuration is shown. The aircraft 100 includes a fuselage 102, wings 104 mounted to the fuselage 102, and one or more rear stabilizers 106 mounted to the rear of the fuselage 102. The aircraft 100 can be a vertical takeoff and landing (VTOL) aircraft and can be a passenger aircraft. A plurality of rotors 112 are mounted to the wings 104 and are configured to provide lift, e.g., for takeoff and landing. A plurality of pivoting propellers 114 are mounted to the wings 104 and are tiltable between a lift configuration and a forward flight configuration as Figure 1 shown, where in the lift configuration, the plurality of pivoting propellers provide a portion of the lift required for vertical takeoff and landing and hovering, and in the forward flight configuration, the plurality of pivoting propellers provide forward thrust for the aircraft 100 for horizontal flight.

[0070] During takeoff and landing, the pusher propellers 114 are tilted to a lift configuration in which the thrust of the pusher propellers is directed upward to provide lift. For forward flight, the pusher propellers 114 are tilted from their lift configuration to their forward flight configuration in which the thrust of the pusher propellers is directed forward to provide forward propulsion. In other words, the pitch of the pusher propellers 114 varies from an angle of tilt in which the pusher propellers provide lift for takeoff and landing (and optionally hovering) to an angle of tilt in which the pusher propellers provide forward thrust for the aircraft 100 for forward flight. The pusher propellers 114 can each be tilted by one or more actuators. The actuators can be electric. Optionally, each pusher propeller 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 pusher propeller, the damper mechanisms being configured to limit the rate of change of the tilt angle of the tiltable pusher propellers, for example in the case where the actuators are disconnected or otherwise fail.

[0071] When the aircraft 100 is in full forward flight, lift can be provided entirely by the wings 104, and the rotors 112 can be shut down. The blades 120 of the rotors 112 can be locked in a low drag position for aircraft cruising. In some embodiments, the rotors 112 can each have two blades 120, the two blades being lockable in a minimum drag position for cruising, in which one blade is directly in front of the other blade, as Figure 1 shown. In some embodiments, the rotors 112 have more than two blades. In some embodiments, the pusher propellers 114 include more blades 116 than the rotors 112. For example, as Figure 1 shown, the rotors 112 can each include two blades, while the pusher propellers 114 can each include five blades. According to various embodiments, the pusher propellers 114 can have 2 to 5 blades.

[0072] According to various embodiments, the aircraft can include only one wing 104 on each side of the fuselage 102 (or a single wing extending across the entire aircraft), at least a portion of the rotors 112 being located behind the wing 104, and at least a portion of the pusher propellers 114 being located in front of the wing 104. In some embodiments, all of the rotors 112 are located behind the wing 104, and all of the pusher propellers are located in front of the wing 104. According to some embodiments, all of the rotors 112 and pusher propellers 114 are mounted to the wings, i.e., no rotors or pusher propellers are mounted to the fuselage. According to various embodiments, all of the rotors 112 are located behind the wing 104, and all of the pusher propellers 114 are located in front of the wing 104. According to some embodiments, all of the rotors 112 and pusher propellers 114 are positioned inside the wing tips 109.

[0073] According to various embodiments, the rotor 112 and the propeller 114 are mounted to the wing 104 via the strut 122. The strut 122 can be mounted below the wing 104, on top of the wing, and / or can be integrated into the wing profile. According to various embodiments, one rotor 112 and one propeller 114 are mounted to each strut 122. The rotor 112 can be mounted at the rear end of the strut 122, and the propeller 114 can be mounted at the front end of the strut 122. In some embodiments, the rotor 112 is mounted in a fixed position on the strut 122. In some embodiments, the propeller 114 is mounted to the front end of the strut 122 via a hinge or other system. The propeller 114 can be mounted to the strut 122 such that when in the forward flight configuration, the propeller 114 is aligned with the body of the strut 122, thereby forming a continuous extension of the front end of the strut 122, which minimizes the drag during forward flight.

[0074] During takeoff and landing, the vehicle is operated by positioning the propellers in a lift configuration and providing the required lift for the vehicle via the combined lift provided by the rotors and the propellers. According to various embodiments, during takeoff and landing and / or hovering, the propellers can be maintained in a predetermined lift configuration, which can be the same for all propellers or different for different propellers. According to various embodiments, the pitch of at least some of the propellers can be actively adjusted during takeoff and landing and / or hovering to provide the required stability and / or maneuverability. As discussed further below, the pitch of the blades of the propeller (also referred to herein as the pitch angle and the angle of attack) is mechanically related to the tilt of the propeller such that the blade pitch is coordinated with the propeller tilt to achieve maximum thrust when the propeller is in the lift configuration and to achieve increased efficiency when the propeller is in the forward flight configuration.

[0075] According to various embodiments, each rotor and / or each propeller can be independently controlled by a flight controller according to various degrees of freedom of operation. According to various embodiments, the only degree of freedom of the rotor is the rotational speed of the rotor. According to various embodiments, the degrees of freedom of at least a portion of the propeller include the rotational speed of the propeller and the degree of tilt of the propeller (in combination with the blade pitch of the propeller). According to various embodiments, any of these degrees of freedom can be actively controlled by the flight controller during takeoff and landing (automatically or in response to a pilot command) to provide appropriate stability and maneuverability.

[0076] Once the aircraft reaches a sufficient altitude to begin forward flight, the proprotor begins to tilt forward towards its forward flight configuration such that its thrust provides a combination of lift and thrust, where the proportion of lift gradually decreases as the proprotor tilts further towards its forward flight configuration. When the proprotor tilts forward towards its forward flight configuration, the pitch angle of the rotor blades can be adjusted. For example, in the forward flight configuration, the rotor blades of the proprotor can be at a pitch angle that results in less drag relative to the pitch angle when the rotor blades are in a lift configuration. The rotor can remain active during at least a portion of the period when the proprotor tilts forward to continue to provide rotor-based lift. The rotor can be deactivated at any time after the forward airspeed is high enough such that the lift provided by the wings is sufficient to maintain the altitude of the aircraft.

[0077] The tilt of at least some of the proprotors can be actively controlled during cruise to provide additional stability and / or maneuverability control. In some embodiments, the tilt of at least some of the proprotors is actively controlled during takeoff and landing and / or hover. In some embodiments, the tilt of the proprotors is fixed (i.e., unchanging) during cruise. According to some embodiments, the tilt of the outermost proprotors can be actively and independently controlled during vertical takeoff and landing and / or hover to provide a yaw moment as needed. The tilt angle range of the tiltable proprotors is at least 90 degrees such that the proprotors can tilt between a forward flight configuration and a lift configuration.

[0078] Figure 2 for an aircraft 100 according to one or more examples of the present disclosure Figure 1 perspective view, which shows the positions of the proprotors in the lift configuration and the forward flight configuration. The proprotor 114 can tilt about a tilt axis 118 that is perpendicular to the forward direction of the aircraft. For forward flight, the proprotor tilts from a lift configuration that provides vertical thrust to a forward flight configuration that provides forward thrust.

[0079] As further described below, the aircraft 100 can include a system that tilts the proprotor 114 between a lift configuration and a forward flight configuration. The system can mechanically correlate the adjustment of the pitch angle of the rotor blades of the proprotor 114 with the tilt of the proprotor 114.

[0080] Figure 3 An exemplary system 202 for coupling tilting of the proprotor with adjusting the pitch angle of the rotor blades of the proprotor is shown. The system 202 is configured such that the pitch angle of the rotor blades of the proprotor can correspond to the tilt position of the proprotor without the need for separate systems for tilting the proprotor and adjusting the pitch angle of the rotor blades. Thus, the system 202 reduces the complexity and cost of the aircraft.

[0081] System 202 rotatably couples the swashplate 203 to the boom 205 of the aircraft and rotatably couples the swashplate 114 to the boom 122 of the aircraft 100. System 202 may include a fixed frame 220 for mounting to the aircraft (e.g., the swashplate 203 may be connected to the boom 205 via a bracket 236 of the fixed frame 220) and a swashplate frame 222 to which the swashplate 203 is mounted. The swashplate frame 222 may be tiltably connected to the fixed frame 220 at the joint 212.

[0082] System 202 may include one or more arms 204 connected to the swashplate frame 222. A linear actuator 206 may be connected to the arm 204 to tilt the swashplate 203 about the 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 a servo motor may be mounted at the joint or may drive a gear train having an output gear located at or engaging a gear located at the joint.

[0083] System 202 may include a cam 214 that may rotate corresponding to the tilt of the swashplate 203. A control rod (discussed further below) may be operatively coupled to the cam 214 such that when the cam 214 rotates, the control rod may translate. The control rod is coupled at its opposite ends to the blades 250 of the swashplate 203 such that translation of the control rod adjusts the pitch angle of the blades 250 of the swashplate 203. Thus, system 202 correlates the pitch angle of the blades 250 of the swashplate 203 with the tilt of the swashplate 203. During operation, the control system of the aircraft may send a swashplate tilt adjustment command to the actuator 206. The actuator 206 may extend or retract such that the swashplate increases or decreases its tilt. As the swashplate tilt changes, the cam 214 rotates. This in turn may cause the control rod to translate, thereby adjusting the pitch of the blades 250.

[0084] Figure 4 A detailed view showing an exemplary system 302 for mechanically correlating tilting of a swashplate of an aircraft between a vertical thrust position and a forward thrust position and adjusting the pitch angle of the blades of the swashplate. System 302 may be used for Figure 3 system 202. System 302 may include a fixed frame 320 for mounting to a portion of a VTOL aircraft (e.g., mounting to a fuselage, wing, or boom structure) and a swashplate frame 322 for mounting a swashplate 303 that is rotatably mounted to the fixed frame 320 about a rotational axis 301. In one or more examples, system 302 includes a gear 308, a pinion (not shown in the figure), a cam 312, a control rod 314, and a pair of arms 304.

[0085] As Figure 4As shown, the gear 308 is positioned along the axis of rotation 301. The gear 308 can be fixed in place relative to the fixed frame 320. For example, as Figure 4 shown, the gear 308 is connected to an internal pin 318 attached to the fixed frame 320. One or more of the shafts 316 can surround the internal pin 318, which is clearly shown in the cross-sectional view of the left shaft 316 of Figure 4 . A set of bearings 317 is located between each shaft 316 and the fixed frame 320 such that the shafts 316 are rotatably mounted to the fixed frame 320.

[0086] The proximal end of the arm 304 can engage with the paddle frame 322. The engagement between the arm 304 and the paddle frame 322 can be a fixed connection, such as by bolting or welding the arm 304 to the paddle frame 322. Optionally, both the arm 304 and the paddle frame 322 can be fixedly connected to the shaft 316. The distal end of the arm 304 can be connected to one or more actuators 306 (see Figure 5 ), which drive the arm 304 to rotate about the axis of rotation 301. When the actuator 306 drives the arm 304, the paddle 303 rotates about the axis of rotation 301.

[0087] In Figure 5 the engagement between the gear 308, the pinion, the cam 312 and the control rod 314 is shown more clearly, which shows Figure 4 a detailed view of a portion of the exemplary system 302. The pinion 310 is mounted within a pinion housing 309, which is coupled to the arm 304 and / or the paddle frame 322 in a fixed position. Thus, the pinion 310 is rotationally coupled to the paddle frame 322 such that the pinion 310 moves with the paddle frame 322 (e.g., as the paddle frame 322 and the paddle 303 rotate about the axis of rotation 301). The pinion 310 also engages with the gear 308. Rotational drive of the pinion housing 309 drives the pinion 310 around at least a portion of the gear 308, which causes the pinion 310 to rotate via engagement with the teeth of the gear 308.

[0088] The cam 312 is fixedly connected to the pinion 310, for example via an internal pin 311, such that the cam 312 rotates with the pinion 310. The cam 312 is also operatively coupled to the first end of the control rod 314 such that the control rod 314 translates relative to the internal pin 311 during at least a portion of the rotation of the cam 312. The control rod 314 is coupled to the blade of the paddle at the second end (as will be described below) such that translation of the control rod 314 adjusts the pitch angle of the blade.

[0089] When the pitch frame 322 rotates about the axis of rotation 301 (e.g., to tilt the pitch propeller 303), the pinion 310 rotates about the gear 308, which causes the cam 312 to rotate and the control rod 314 to translate, thereby adjusting the pitch angle of the blades of the pitch propeller 303. Thus, the system 302 mechanically couples tilting the pitch propeller 303 with adjusting the pitch angle of the blades of the pitch propeller 303.

[0090] The control rod 314 can engage the cam 312 via a follower that follows the cam 312 as the cam 312 rotates. The follower can be, for example, a roller or a pin. As Figure 5 shown, the control rod 314 engages the cam 312 via the roller 313. As the cam 312 rotates, the roller 313 travels along the outer surface of the cam 312. To maintain engagement with the outer surface of the cam 312, the control rod 314 can be biased against the cam 312 in a compressed manner, such as via a spring (not shown in the figure).

[0091] The cam profile (e.g., the shape of the outer surface of the cam followed by the control rod) controls the position of the control rod. The profile of the cam 312 can include one or more portions that cause the control rod 314 to translate and can include one or more portions that do not cause the control rod 314 to translate. For example, the cam 312 can have one or more helical portions that cause the control rod 314 to translate and / or one or more circular portions that do not cause the control rod 314 to translate. In Figure 5 the example shown, the cam 312 includes a helical profile that will cause the control rod 314 to translate continuously throughout the tilt range of the pitch propeller.

[0092] Figure 6 Show Figure 4 and Figure 5 Another detailed view of a portion of the exemplary system 302. Although Figure 5 depicts the control rod 314 engaging both sides of the outer surface of the cam 312 (e.g., in the corner region of a 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 obtained by rotating the cam 312 in the clockwise direction such that the roller 313 moves in the counterclockwise direction as it follows the surface of the cam 312.

[0093] When roller 313 follows the helical portion of cam 312, control lever 314 can translate towards or away from internal pin 311 at the center of cam 312. For example, if cam 312 rotates in a clockwise direction, then as roller 313 follows the helical portion of cam 312, control lever 314 translates away from internal pin 311. Conversely, if cam 312 rotates in a counterclockwise direction, then as roller 313 follows the helical portion of cam 312, control lever 314 can translate towards internal pin 311. When roller 313 follows the circular portion of cam 312, control lever 314 can remain at a constant distance and not translate relative to internal pin 311. For example, cam 312 can include a helical profile for the first 210 degrees of rotation and a circular profile for the remaining 150 degrees of rotation such that control lever 314 only translates away from internal pin 311 during the first 210 degrees of rotation of cam 312. As mentioned above, control lever 314 can be biased against cam 312 in a compressed manner via spring 315.

[0094] Figure 6 An example showing that control lever 314 has a roller 313 that engages the outer surface of cam 312. Alternative configurations can have different engagements between the control lever and the cam. For example, Figure 7 A top perspective view of an exemplary system 402 including a control lever 414 having a pin 413 that engages a track 411 of a cam 412. Pin 413 engages track 411 of cam 412 such that cam 412 can push and pull pin 413 and the control lever as cam 412 rotates.

[0095] System 402 can be used for Figure 3 system 202. Similar to the system discussed above, system 402 connects a paddle 403 to a part of the aircraft (connected to a boom 401 as Figure 7 shown). However, different from the above system, system 402 includes a control lever 414 having a pin 413 that is located in a track. This configuration can more clearly show the pop - out details of cam 412, which shows the engagement of pin 413 of control lever 414 with track 411 in cam 412. When cam 412 rotates, for example, by engagement with a pinion 410 that rotates around a gear 408, pin 413 can follow track 411, thereby translating control lever 414.

[0096] Gear 408 can be fixed in place relative to a fixed frame 420 that is fixedly mounted to the aircraft. For example, as Figure 7 shown, gear 408 is connected to fixed frame 420. Pinion 410 can be rotatably mounted to fixed frame 420 such that pinion 410 is rotationally coupled to a paddle frame 422 and moves with paddle frame 422.

[0097] The track 411 may include a helical portion and a circular portion. When the pin 413 follows the helical portion of the track 411, the control lever 414 may translate towards or away from the center of the cam 412. However, when the pin 413 follows the circular section of the track 411, the control lever 414 may remain at a constant distance and not translate relative to the center of the cam 412. Optionally, based on the desired type of translation, the track 411 of the cam 412 may be of various geometries. To maintain engagement with the track 411, the control lever 414 may be biased against the cam 312, for example, via a spring (not shown in the figure) in a compressed or stretched manner.

[0098] Figure 8 Another exemplary cam control lever configuration is shown. Different from the system 302, in the system 502, the control lever 514 is kept in a stretched state. Similar to the above system, the system 502 may include a paddle frame 522 rotatably mounted to a fixed frame 520, and the fixed frame is mounted to a part of the aircraft (such as a fuselage, a wing, or a boom structure). The system 502 may be configured such that one or more actuators (not shown in the figure) drive the paddle frame 522 to rotate about the axis of rotation 501 to tilt the paddle mounted to the paddle frame 522 between a vertical thrust position and a forward thrust position.

[0099] Similar to the above system, the system 502 includes a control lever 514 that engages with a cam 512, and the cam rotates based on engagement with a pinion 510 that engages with a gear 508. The gear 508 may be fixed in place relative to the fixed frame 520. For example, as Figure 7 shown, the gear 508 is connected to the fixed frame 520. The cam 512 is fixedly connected to the pinion 510 via a shaft 516, such that the cam 512 rotates with the pinion 510. The pinion 510 may be rotatably mounted to the fixed frame 520, for example, via a bearing mounted to a rib 560, such that the pinion 510 is rotationally coupled to the paddle frame 522 and moves with the paddle frame 522. When the paddle frame 522 moves, the pinion 510 rotates around the gear 508, which causes the cam 512 to rotate.

[0100] The control lever 514 includes a U-shaped clip 515 and a follower 513, which is a roller in this example. As shown, the cam 512 engages with the follower 513 such that the follower 513 rolls along the cam as the cam 512 rotates. The follower 513 is rotatably attached to the U-shaped clip 515 of the control lever 514. The control lever 514 is in a stretched state (for example, via one or more springs along the axis 514 towards Figure 8a force is applied to the left side in the view of), such that the follower 513 is forced against the cam 512. Thus, as the follower 513 rolls along the cam 512, the control rod 514 can translate along the axis 505 (depending on the profile of the cam 512). As described above, the cam 512 can have any suitable profile for achieving the desired relationship between blade pitch and swashplate inclination.

[0101] The U-bolt 515 includes a slot 536 through which the shaft 516 can extend. The U-bolt 515 is separated from the shaft 516 by a bushing, which is more clearly shown in Figure 9 as shown. Figure 9 shown Figure 8 a detailed cross-sectional view of the interface of the bushing 519 and the U-bolt 515 of the system 502 shown. The bushing 519 can include a mating surface 529 on the region of the bushing 519 that engages the U-bolt 515, where the mating surface 529 is shaped such that rotation of the U-bolt 515 is prevented, which in turn prevents rotation of the control rod (e.g., Figure 8 the control rod 514 shown in) attached to the U-bolt 515. As shown in Figure 9 shown, the mating surface 529 of the bushing 519 is flat, corresponding to the flat surface of the U-bolt 515. Optionally, based on the shape of the corresponding surface of the U-bolt 515, these surfaces can be another shape. For example, the engagement between the U-bolt 515 and the bushing 519 can involve circular, elliptical, or angled surfaces.

[0102] As described above, the translation of the control rod of any of the above exemplary systems can adjust the pitch angle of the blades of the swashplate. The control rod can be operably engaged with a plurality of blades such that translation of the control rod causes rotation of the blades. Figure 10 A detailed cross-sectional view of an exemplary hub 1000 of a swashplate according to one or more examples of the present disclosure is shown, showing an example of the connection of the control rod 1014 to the blades 1012 of the swashplate. The hub 1000 can include an engine shaft 1004, a spring 1015, a bearing 1003, a plate 1002, a pitch plate 1005, and a plurality of linkages 1006. The blades 1012 are mounted to the hub 1000. For example, each blade 1012 can include a blade root 1010 connected to the linkage 1006. The control rod 1014 is connected to the pitch plate 1005 and / or the plate 1002 of the hub 1000. As described above, the other end of the control rod 1014 can be connected to a system configured to translate the control rod 1014 while tilting the swashplate.

[0103] The engine shaft 1004 is disposed about the spring 1015 and the bearing 1003 and is connected to the plate 1002, which is connected to the pitch plate 1005 that engages the link 1006. The engine shaft 1004 is connected to an engine (not shown in the figure) of the hub 1000. The plate 1002 is rotationally restricted by a spline interface with the engine shaft 1004 such that the plate 1002 rotates with the engine shaft 1004. The control lever 1014 is prevented from rotating with the plate 1002 via the bearing 1003. The pitch plate 1005 is connected to the blade root 1010 via the link 1006. As Figure 10 shown, the link 1006 is a dog-bone link, although other types of link mechanisms, such as pitch links, etc., are encompassed. The spring 1015 can maintain the control lever 1014 in a stretched or compressed state. In Figure 10 the example of, the spring 1015 holds the control lever 1014 in a stretched state by pressing the control lever 1014 against the pitch plate 1005.

[0104] When the control lever 1014 translates (e.g., axially advances or retracts), the plate 1002 and / or the pitch plate 1005 translate, which causes the link 1006 to adjust the pitch angle of the blade 1012 by rotating each blade 1012 about the central axis 1008 of the blade 1012. Rotating each blade 1012 about the central axis 1008 adjusts the pitch angle of the blade 1012. Thus, translation of the control lever 1014 can adjust the pitch angle of the blades 1012 of the variable pitch propeller.

[0105] As described above, adjusting the pitch of the blades of a variable pitch propeller based on the inclination of the variable pitch propeller can enable tuning of the propeller operating characteristics for different flight phases, which can improve efficiency and thus reduce the energy requirements during flight. The relationship between the blade pitch and the inclination of the variable pitch propeller can be selected by choosing a desired cam profile (e.g., the profile of the cam 214). Based on the selection of the cam profile, various relationships can be achieved between the blade pitch and the inclination of the variable pitch propeller. Figure 11 is a graph showing examples of some of these relationships.

[0106] Figure 11 The graph of shows the blade pitch varying with the variable pitch propeller tilt angle. In Figure 11 the variable pitch propeller tilt angle provided on the X-axis of the graph is the angle formed between the axis of rotation of the variable pitch propeller and a line extending parallel to the longitudinal axis of the aircraft and intersecting the axis of rotation of the variable pitch propeller. When the variable pitch propeller is in the forward flight position, the axis of rotation of the variable pitch propeller is Figure 2 shown in the example of as the axis of rotation 1101-A (the axis of rotation 1101-A coincides with the line extending parallel to the longitudinal axis 1102 of the aircraft and intersecting the axis of rotation of the variable pitch propeller), and when the variable pitch propeller is in the lift position, it is shown as the axis of rotation 1101-B. Figure 11The zero tilt angle on the curve graph corresponds to the rotational axis of the pitch propeller being parallel to the longitudinal axis 1102 of the aircraft - for example, for providing forward thrust for forward flight, while the ninety-degree tilt angle corresponds to the pitch propeller providing vertical thrust, such as for vertical takeoff and landing.

[0107] Figure 11 The pitch of the blade provided on the Y-axis of the curve graph can be defined as the angle between the chord of the blade and the plane of rotation, and can be measured at a specific point along the length of the blade. Figure 12 A diagram showing an example of the definition of blade pitch. The blade 1208 (only one is shown for simplicity, but it should be understood that each pitch propeller will contain multiple blades) rotates around the pitch propeller shaft 1214 and is mounted such that it can rotate around the pitch axis 1206, thereby enabling the adjustment of the pitch 1216 of the blade 1208. The plane that contains the pitch axis 1206 and is traversed by the blade 1208 during the rotation of the blade 1208 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 connecting the leading edge and the trailing edge of the blade 1208) and the line 1202 perpendicular to the pitch axis 1206 and located within the disk plane 1202. Figure 11 The pitch values shown in [the figure] are merely exemplary, and those of ordinary skill in the art will understand 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, Figure 11 the pitch propeller tilt values shown in [the figure] are merely exemplary, and those of ordinary skill in the art will understand that different ranges of pitch propeller inclinations can be used, including negative tilt angles associated with slightly downward tilting of the pitch propeller.

[0108] Figure 11 Contains four different lines 1102 - 1108, which indicate four different relationships between blade pitch and pitch propeller inclination. Each line is associated with a different cam profile. Lines 1102, 1104, and 1106 have the maximum blade pitch (the maximum value of the given line) at a zero pitch propeller tilt angle. This can be the blade pitch for forward flight, where the larger the blade pitch, the greater the efficiency provided during forward flight at relatively high airspeeds.

[0109] Line 1108 has a minimal blade pitch at zero swashplate inclination. This may result in a relatively high drag of the blade at high airspeeds, which can be used to decelerate the aircraft, for example, for landing. This small blade pitch can also be used to increase the efficiency of the swashplate at low airspeeds (e.g., during a conventional (airplane type) takeoff). The blade pitch rapidly increases to a maximum value such that the swashplate only needs to tilt by a relatively small amount to achieve the maximum blade pitch, which can be the desired blade pitch for forward flight. Using this relationship, the swashplate only needs to tilt by a small amount relative to the minimum inclination (e.g., zero inclination, small positive inclination, small negative inclination, etc. as shown in the figure) to achieve a better blade pitch for high-speed forward flight.

[0110] Each of lines 1102 - 1108 shows the blade pitch decreasing to a minimum blade pitch (the minimum value for a given relationship), which is associated with the maximum swashplate inclination (the inclination that can be used for vertical flight and hovering). The minimum blade pitch may be optimal for the low airspeeds and high thrust requirements of vertical flight and hovering. Lines 1104 and 1106 show that the minimum blade pitch does not need to be zero. The specific swashplate tilt angle to reach the minimum blade pitch can be selected based on the cam profile, as shown by the different positions of this point (e.g., point 1110) for various relationships.

[0111] Each relationship 1102 - 1108 is achieved through a different cam profile. The cam profile may include multiple regions with different shapes to achieve variations in the relationship between the blade pitch and the swashplate inclination for the various lines 1102 - 1108. For example, line 1102 can be achieved through a cam profile that includes a helical section corresponding to the range of swashplate inclination angles from zero degrees to the swashplate tilt angle at point 1110. The helical section (continuously varying radius) transitions to a circular section (constant radius), which provides a constant blade pitch associated with the swashplate tilt angle passing through point 1110. Line 1106 can be achieved through a first helical section, followed by a second helical section with a different rate of radius change from the first helical section, followed by a circular section. Line 1104 can be achieved through a cam profile with a continuously varying rate of radius change.

[0112] Figure 11 The relationships shown are only examples to illustrate that many different relationships can be achieved through appropriate selection of cam profiles. A person of ordinary skill in the art will understand that the desired relationship between the blade pitch and the swashplate inclination can be achieved through appropriate cam profile design.

[0113] Secondary load path with a passive damper for a tiltable assembly

[0114] As described above, disconnecting the actuator from the tiltable aircraft assembly creates a risk of catastrophic failure caused by uncontrolled movement of the tiltable assembly. The passive damper described herein provides a mechanism for suppressing the high-speed tilting of the tiltable propeller in the event of actuator disconnect. The likelihood of catastrophic failure caused by actuator disconnect can potentially be reduced by: introducing redundant actuators such that if the first actuator fails (e.g., due to breakage of the actuator shaft), then another actuator takes over; and / or over-sizing the aircraft structural components such that a particular failure mode is not catastrophic, but these mitigation methods can be expensive and introduce unwanted additional weight into the aircraft design. The passive damper described herein provides a smaller and cost-effective method of preventing catastrophic failure.

[0115] The passive damper is configured to apply a force (hydraulic or pneumatic) to the connecting component only when the connecting component is in motion (i.e., no bias is applied when the connecting component is stationary). The passive damper described herein can be connected between the strut of the aircraft and the tiltable propeller to apply a damping force when adjusting the tilt angle of the tiltable propeller. As described above, the passive damper can include a balanced hydraulic cylinder or pneumatic cylinder (“balanced non-resistive damper”). The balanced non-resistive damper cylinder can include a piston that slides within the cylinder, where the piston includes a plurality of holes through which damping fluid flows as the piston moves through the cylinder. The piston can be attached to a piston rod extending from the cylinder and can be rotatably attached to the tiltable propeller of the aircraft. When the tiltable propeller rotates towards the lift position or the forward flight position, it will apply a tensile or compressive force to the piston rod depending on the tilt direction, thereby pushing the piston rod away from or towards the piston cylinder. In turn, when the propeller tilts in either direction, the piston will be forced to slide within the cylinder. As the piston moves through the damping fluid, the damper dissipates energy (e.g., by converting kinetic energy into heat) due to the resistance of the damping oil applied to the piston, thereby suppressing the tilt speed of the tiltable propeller.

[0116] In some examples, the passive damper may include an unbalanced cylinder (“unbalanced damper”) in place of a balanced hydraulic or pneumatic cylinder. The unbalanced damper similarly includes a piston that slides within the cylinder and includes a plurality of holes through which damping fluid flows as the piston moves through the cylinder, but in addition to the damping fluid, the cylinder may also contain pressurized gas. When a force is applied to the damper piston, the pressurized gas may compress or expand, depending on the direction of the force. Thus, the pressurized gas can absorb the initial shock force applied to the damper piston. Accordingly, if the actuator shaft breaks, the damper will inhibit the high-speed tilting of the tiltable propeller, thereby preventing a catastrophic failure. Additionally, the passive damper does not require activation, engagement, or other control to perform its function of inhibiting the high-speed tilting of the tiltable propeller. Thus, it provides a reliable and simple solution for the actuator disconnection fault mode described herein.

[0117] In one or more examples, any of the above systems, such as system 202, system 302, system 402, system 502, and aircraft 100, may include a damper as shown in exemplary system 1302 of Figure 13 to limit the rate of tilt of the propeller, for example, in the event of an actuator failure. System 1302 may be implemented in an aircraft having a tiltable propeller, in place of redundant actuators and / or a joint assembly including a fail-safe latch as described below, or in addition to redundant actuators and / or a joint assembly including a fail-safe latch as described below.

[0118] Damper 1325 may include a balanced hydraulic cylinder or a balanced pneumatic cylinder 1316, respectively, as is known in the art, the balanced hydraulic cylinder or balanced pneumatic cylinder including a piston 1318 that slides within the cylinder and is connected to a piston rod 1307 or 1407, where the piston includes a plurality of holes (not shown) through which fluid flows as the piston moves within the cylinder. Damper 1325 of system 1302 is connected between a boom 1305 of an aircraft (such as Figure 1 and 2 aircraft 100) and a propeller frame 1322, where a propeller (not shown in the figure) is mounted to the propeller frame. Damper 1325 may be housed within a housing of boom 1305. In one or more examples, boom 1305 may include ribs 1350, and actuator 1306 and damper 1325 may be positioned on opposite sides of ribs 1350. Propeller frame 1322 is tiltably connected to a fixed frame 1320 such that propeller frame 1322 (and the propeller) can tilt about a rotational axis 1301. Actuator 1306 is connected to propeller frame 1322 and drives the propeller to tilt about rotational axis 1301.

[0119] The damper 1325 may be rotatably attached to the boom 1305 by one or more mounts 1314 on a first side of the rib 1350 and on a second side of the housing of the boom 1305. The damper 1325 is rotatably connected to the swashplate frame 1322 at a connector 1340 of the 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 extending between two sides of the U-bracket 1309. Thus, when the actuator 1306 tilts the swashplate frame 1322 toward the lift position (clockwise with respect to Figure 13 the view of Figure 13 ), the swashplate frame 1322 exerts a tensile force on the piston rod 1307 of the damper 1325, causing the piston rod 1307 to extend out of the cylinder. The fluid within the cylinder resists such movement, thereby generating a reaction force acting on the piston 1318 within the damper 1325, which reaction force is proportional to the speed of movement of the piston. A plurality of holes in the piston allow the damping fluid to flow through the piston 1314, such that the piston can move within the cylinder and, in turn, allow the piston rod 1307 to extend from the cylinder until the swashplate reaches the lift position. Similarly, when the actuator 1306 tilts the swashplate frame 1322 toward the forward flight position (counterclockwise with respect to

[0120] the view of Figure 14 ), the swashplate frame 1322 exerts a compressive force on the piston rod 1307 of the damper 1325, causing the piston rod 1307 to retract into the cylinder 1316. The fluid within the cylinder resists such movement in a similar manner, thereby generating a reaction force acting on the piston 1318 within the damper 1325, which reaction force is proportional to the speed of movement of the piston. Figure 13 In one or more examples, the actuator 1305 may be a linear actuator. Alternatively, the actuator may be a rotary actuator, as shown in the exemplary system 1402 of

[0121] The system 1402 is otherwise similar to Figure 14If the actuators 1306 or 1406 are disconnected from the swashplate frames 1322 or 1422 respectively, a catastrophic failure may occur because the swashplate may start to tilt rapidly out of control. In one or more examples, the system 1302 or 1402 may include a second (redundant) actuator that is also connected to the swashplate frame 1322 or 1422 such that if the first actuator 1306 or 1406 is disconnected, the second actuator may still control the tilt of the swashplate about the axis of rotation 1301 or 1401. However, adding a second actuator may complicate the system 1302 or 1402, increase cost, and add more weight to the aircraft. Instead of implementing a second actuator, the systems 1302 and 1402 actually include damper mechanisms, such as dampers 1325 or 1425. In the case where the actuators 1306 or 1406 are disconnected from the swashplate frames 1322 or 1422, the dampers 1325 or 1425 dissipate energy, thereby suppressing the tilt speed of the swashplate, which eliminates the catastrophic consequences of the disconnection of the actuators 1306 or 1406 from the swashplate frames 1322 or 1422.

[0122] As described above, the damper 1325 or 1425 may be rotatably connected to a pin 1308 or 1408 that extends between the two sides of the clevis bracket 1309 or 1409, and the clevis bracket is fixedly attached to the swashplate frame 1322 or 1422 respectively. If the actuator 1306 (or Figure 14If the actuators 1406 are disconnected from the swashplate frames 1322 or 1422 respectively, then the dampers 1325 or 1425 will inhibit the tilting speed of the swashplate during swashplate movement, but will not apply a damping force when the swashplate is stationary. When the swashplate tilts towards the forward flight position, the swashplate frames 1322 or 1422 apply a compressive force to the damper piston rods 1307 or 1407 via the U-bolt brackets 1309 or 1409, thereby causing the damper piston rods 1307 or 1407 to retract into the cylinders 1316 or 1416. Similarly, when the swashplate frames 1322 or 1422 (and the swashplate) tilt towards the lift position, the swashplate frames 1322 or 1422 apply a tensile force to the damper piston rods 1307 or 1407 via the U-bolt brackets 1309 or 1409, thereby causing the piston rods 1307 or 1407 to extend outward from the piston cylinders 1316 or 1416. When the swashplate frames 1322 or 1422 (and the swashplate) tilt in either direction after the actuators are disconnected, the fluid in the cylinder resists the extension or retraction of the damper piston rods 1307 or 1407 by applying a reaction force acting on the pistons 1318 or 1418 within the dampers 1325, and the reaction force is proportional to the moving speed of the pistons. Multiple holes in the pistons allow the damping fluid to flow through the pistons, enabling the pistons to move within the cylinders and, in turn, allowing the damper piston rods 1307 or 1407 to extend or retract at a damped speed. Thus, the dampers 1325 or 1425 dissipate the rotational energy of the swashplate by applying a force in a direction opposite to the tilting direction of the swashplate.

[0123] As mentioned above with respect to systems 1302 and 1402, dampers 1325 and / or 1425 may each include a balanced hydraulic cylinder or a balanced pneumatic cylinder 1316 or 1416, known in the art, which includes a piston 1318 or 1418 that slides within the cylinder and is connected to a piston rod 1307 or 1407, where the piston includes a plurality of holes (not shown) through which fluid flows as the piston moves within the cylinder. A balanced damper contains less damping fluid than an unbalanced damper and is thus generally less massive than an unbalanced damper. The balanced dampers disclosed herein may require from 0.25 to 0.5 quarts of damping fluid. In some examples, the balanced dampers require less than 0.35 quarts of damping fluid. In contrast, conventional dampers may require from 0.51 quarts to 1.0 quarts of damping fluid. In some examples, the conventional dampers disclosed herein may require at least 0.74 quarts of damping fluid. In some examples, damper 1325 or 1425 may include a fire-retardant damping fluid, such as MIL-PRF-83282 or MIL-PRF-87257. MIL-PRF-83282 is a fire-retardant damping fluid; however, the fluid becomes more viscous at low temperatures, which causes more parasitic drag during low-temperature operation. MIL-PRF-87257 maintains a similar fire-retardancy to MIL-PRF-83282 but has improved low-temperature viscosity properties.

[0124] As mentioned above, dampers 1325 and 1425 are passive dampers that are configured to apply a force (hydraulic or pneumatic) to the swashplate frame 1322 or 1422 (and the swashplate) only when the tilt angle of the swashplate is being adjusted (i.e., no bias is applied when the swashplate is stationary). In one or more examples, systems 1302 and 1402 may each include one or more redundant actuators and a damper mechanism such as damper 1325 or 1425.

[0125] Damper 1325 or 1425 may be configured to limit the rate of change of the tilt angle of swashplate frame 1322 or 1422 in two tilt directions, respectively. Damper 1325 or 1425 may be positioned such that the force vector of damper 1325 or 1425 may be along the tilt axis (e.g., Figure 2extends below the tilt axis 118). In one or more examples, the damper 1325 or 1425 can be configured to limit the rate of change of the tilt angle of the swashplate in the case where the actuator 1306 or 1406 is disconnected from the swashplate frame 1322 or 1422, respectively. According to some examples, the dampers described herein can limit the rate of change of the swashplate inclination to a maximum of 100 degrees per second under maximum operating load. In some examples, the damper can limit the rate of change of the swashplate tilt angle between 100 degrees per second and 300 degrees per second. In some examples, the damper can limit the rate of change of the swashplate tilt angle between 100 degrees per second and 200 degrees per second. In some examples, the damper can limit the rate of change of the swashplate tilt angle between 150 degrees per second and 180 degrees per second. In some examples, the damper can limit the rate of change of the swashplate 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.

[0126] As described above, in some examples, the damper 1325 or 1425 can be a balanced non-resistive damper. In such examples, the damper 1325 or 1425 can be configured to handle peak damping pressures between 2000 psi and 4000 psi. In some examples, the damper 1325 or 1425 can be configured to handle a peak damping pressure of approximately 3000 psi. The damper 1325 or 1425 can be configured with an outer cylinder diameter between one inch and two inches, a piston diameter between 0.25 inches and 0.75 inches, a stroke between nine inches and 11 inches, a body length between 10 inches and 12 inches, and an orifice diameter between 0.04 inches and 0.07 inches. The mass of the damper 1325 or 1425 can be between two pounds and four pounds. For example, an exemplary balanced non-resistive damper weighs approximately 2.2 pounds, and this weight can include uncertainties for the damper cylinder, damper rod / piston, and damper fluid, as well as bearings, rod ends, trunnions, and various other accessories. The peak power consumed by the damper during the transition (i.e., during normal operation when the swashplate transitions between the forward flight configuration and the lift / hover configuration) can be 0.7 watts, while the peak power consumed by the damper during hover (i.e., during normal operation when stabilizing the swashplate tilt angle in the lift / hover configuration) can be 6.3 watts.

[0127] In other examples, the damper 1325 or 1425 can be an imbalance damper. In such examples, the damper 1325 or 1425 can be configured to handle a maximum damping pressure between 2000 psi and 4000 psi, and in some examples can be configured to handle a maximum damping pressure of 3000 psi. The imbalance damper can be configured to operate at a working pressure (or pre-charge) between 1000 psi and 2000 psi, and in some examples, the working pressure / pre-charge can be configured to be approximately 1500 psi. The loads from different regions of the imbalance damper 1325 can be between 300 lbf and 400 lbf, and in some examples, the loads from different regions of the imbalance damper 1325 or 1425 can be approximately 330 lbf. The damper 1325 or 1425 can be configured with an outer cylinder diameter between 1.5 inches and 2.5 inches, an inner rod diameter between 0.25 inches and 0.75 inches, a stroke between 9 inches and 11 inches, a body length between 13 inches and 14 inches, and an orifice diameter between 0.08 inches and 0.10 inches. The mass of the damper 1325 or 1425 can be between two pounds and four pounds. For example, an exemplary imbalance damper weighs approximately 2.9 pounds, and this weight can include the damper cylinder, the damper rod / piston, and the damper fluid, as well as uncertainties for bearings, rod ends, trunnions, and various other accessories. The maximum power consumed by the damper 1325 or 1425 during transition can be between 25 watts and 30 watts, while the peak power consumed by the damper 1325 during hover can be between 60 watts and 65 watts.

[0128] According to the corresponding exemplary design specifications for a balanced undamped and an unbalanced damped device as described above, the balanced undamped device can provide a favorable configuration with respect to power consumption and can be of smaller mass than the unbalanced damped device, mainly due to its lack of a precharge. Without departing from the scope of the claims, various additional or different design modifications can be made to the exemplary dampers described herein (e.g., dampers 1325 and 1425 described above, and dampers 1525 and 1625 described below). For example, a passive damper can be configured for stroke-sensitive damping, which can be achieved using one or more axial grooves in a portion of the damper cylinder with reduced stroke, or using a stepped hole / piston with a secondary flow path having additional restrictions. The passive damper can also be configured with a stepped valve tuning, an increasing rate shim stack, and / or a pressure release in the stepped valve tuning to allow a wider range of rates approaching the design operating load. Additionally, the passive damper can include an electromagnetic directional valve with a pressure release device. The passive damper can also include various design features to mitigate potential failures of the damper. For example, the damper can include a sight glass for inspecting the oil, or a pressure sensor configured to monitor and verify the nominal damping at the normal operating rate towards an end stop of the aircraft.

[0129] In one or more examples, controlling an aircraft that includes one or more dampers as described above can include: receiving, at a controller, a command to adjust an inclination angle of a tiltrotor that can tilt between a lift position for providing lift to the aircraft and a forward flight position for providing forward propulsion to the aircraft; and controlling at least one actuator in accordance with the command to adjust the inclination angle of the tiltrotor, wherein at least one passive damper is connected to the tiltrotor to limit a rate of change of the inclination angle of the tiltrotor.

[0130] Figure 15A and Figure 15B illustrates a system 1502 similar to systems 1302 and 1402 shown above in Figure 13 and Figure 14 and that is mounted between a pylon of the aircraft and the tiltrotor. The tiltrotor (omitted from Figure 13 and Figure 14 ) is shown in Figure 15A in a lift position and in Figure 15B in a forward flight position. Thus, Figure 15A and Figure 15B depict how the damper 1525 moves in response to a change in the inclination angle of the tiltrotor.

[0131] Figure 15A depicts a first configuration of the system 1502 that includes the damper 1525. The system 1502 includes a rotor 1503 (which is shown inFigure 13 (omitted), the paddle is in a lift configuration and is rotatably connected to a linear damper 1525 and a linear actuator 1506. The damper 1525 and the actuator 1506 of the system 1502 are connected between, for example, Figure 1 and Figure 2 a boom 1505 of an aircraft 100 of an aircraft such as

[0132] and the paddle frame 1522, where the paddle 1503 is mounted to the paddle frame. In one or more examples, the boom 1505 may include ribs 1550, and the actuator 1506 and the damper 1525 may be positioned on opposite sides of the rib 1550. The boom 1505 may also include one or more partitions 1555, and a portion of the piston rod 1507 of the damper 1525 extends through the partition. The paddle frame 1522 may be tiltably connected to a fixed frame 1520 such that the paddle frame 1522 (and the paddle 1503) can be tilted about a rotational axis 1501. The actuator 1506 may be connected to the paddle frame 1522 and drive the paddle to tilt about the rotational axis 1501. In one or more examples, the actuator 1505 may be a linear actuator.

[0132] The actuator 1506 may be configured to tilt the paddle 1503 between Figure 15A the lift configuration depicted in Figure 15B and Figure 15A the forward flight configuration depicted in Figure 15B . When the actuator tilts the paddle 1503 between Figure 15A the lift configuration depicted in Figure 15B and Figure 13 the forward flight configuration depicted in Figure 14 , the damper 1525 may apply a force in a direction opposite to the movement of the paddle. When the actuator tilts the paddle 1503, the fluid in the cylinder 1516 of the damper 1525 is used to resist the extension or retraction of the damper piston rod 1507 by applying a reaction force acting on a piston (not shown) in the cylinder 1516 of the damper 1525, and the reaction force is proportional to the moving speed of the piston. A plurality of holes (not shown) in the piston allow the damping fluid to flow through the piston such that the piston can move within the cylinder. In turn, the piston rod 1507 of the damper 1525 can extend or retract, thereby allowing the paddle to tilt under the influence of the actuator during normal operation. Similar to Figure 13 the damper 1325 of Figure 14 and

[0133] As described above, in the absence of damper 1525, if actuator 1506 becomes disconnected from swashplate frame 1522 (e.g., due to breakage of the actuator shaft), swashplate 1503 may rapidly tilt in an uncontrolled manner, ultimately resulting in a catastrophic failure of the aircraft. Damper 1525 is configured to control the rate of change of the tilt angle of swashplate 1503 such that if actuator 1506 becomes disconnected from the swashplate frame, swashplate 1503 is prevented from tilting at a rate that could damage the aircraft. Damper 1525 is configured to apply a compressive or tensile force to swashplate frame 1522 in accordance with the tilt direction in substantially the same manner as when the actuator is connected to the swashplate, thereby suppressing the tilt speed of the swashplate when the actuator becomes disconnected.

[0134] Figure 16 System 1602 is shown, which includes a damper in an alternative configuration to the configuration shown above in Figure 13 , Figure 14 and Figures 15A - 15B In one or more examples, any of the above systems, such as systems 202, 302, 402, 502, and aircraft 100, may include a damper, as shown in exemplary system 1602 of Figure 16 for example, to limit the tilt rate of the swashplate in the event of an actuator failure. System 1602 may be implemented in an aircraft having a tiltable swashplate, in place of redundant actuators and / or a joint assembly including a fail-safe latch as described below or in addition to redundant actuators and / or a joint assembly including a fail-safe latch as described below.

[0135] Damper 1625 of system 1602 is connected between a fixed frame 1620 of a boom 1605 of an aircraft (e.g., aircraft 100 of Figure 1 and Figure 2 ) and swashplate frame 1622, to which a swashplate (not shown in the figure) is mounted. Actuator 1606 may also be connected between fixed frame 1620 and swashplate frame 1622 and drives the swashplate to tilt about a rotational axis 1601. In one or more examples, actuator 1605 may be a linear actuator. Both damper 1625 and actuator 1606 may be rotatably connected to fixed frame 1620 of boom 1605 and rotatably connected to swashplate frame 1622 such that the swashplate frame (and the swashplate) may rotate relative to damper 1625 and actuator 1606.

[0136] Similar to dampers 1325 and 1425, if actuator 1606 becomes disconnected from swashplate frame 1622 (e.g., due to a break in the actuator shaft), a catastrophic failure can occur because the swashplate may begin to tilt rapidly and uncontrollably. In one or more examples, system 1602 can include a second (redundant) actuator that is also connected to swashplate frame 1622 such that if first actuator 1606 becomes disconnected, the second actuator can still control the tilt of the swashplate about axis of rotation 1601. However, as mentioned throughout, adding a second actuator can complicate system 1602, increase cost, and add more weight to the aircraft. Instead of implementing a second actuator, system 1602 actually includes a damper mechanism, such as damper 1625. In the event that actuator 1606 becomes disconnected from swashplate frame 1622, damper 1625 dissipates energy, thereby limiting the rate of change of tilt of the swashplate (e.g., by applying a force to the damper piston via a damping fluid as described above), which eliminates the catastrophic consequences of actuator 1606 becoming disconnected from swashplate frame 1622.

[0137] Similar to dampers 1325, 1425, and 1525, damper 1625 can be rotatably connected to a pin (not shown) that extends between two sides of a U-bracket 1609 that is fixedly attached to swashplate frame 1622. If actuator 1606 becomes disconnected from swashplate frame 1622, then as swashplate frame 1622 (and the swashplate) tilts towards the forward flight position, damper 1625 will apply a compressive force to swashplate frame 1622 via U-bracket 1609, and as swashplate frame 1622 (and the swashplate) tilts towards the lift position, the damper will apply a tensile force to swashplate frame 1622 via the U-bracket 1609. Thus, damper 1625 dissipates energy by applying a force in a direction opposite to the tilt direction of the swashplate.

[0138] Damper 1625 can be configured to handle a maximum operating pressure between 2000 psi and 4000 psi. In some examples, the maximum operating pressure of damper 1625 can be about 3000 psi. The outer diameter of the cylinder 1616 of damper 1625 can be between 1.0 inch and 2.5 inches. In some examples, the outer diameter of damper 1625 can be about 2.0 inches. The piston diameter 1607 of damper 1625 can be between 0.25 inch and 0.75 inch. In some examples, the diameter of piston 1607 can be about 0.5 inch. The normal operating stroke of damper 1625 can be between 3 inches and 6 inches. In some examples, the normal operating stroke of damper 1625 can be about 4.65 inches. The pin-to-pin range of the normal operating stroke of damper 1625 can be between 9 inches and 16 inches. In some examples, the pin-to-pin range of the normal operating stroke of damper 1625 can be between 9.9 inches and 14.6 inches. The mechanical stroke of damper 1625 can be between 3 inches and 6 inches. In some examples, the mechanical stroke of damper 1625 can be about 5.06 inches as the minimum value. The pin-to-pin range of the mechanical stroke can be between 9 inches and 15 inches. In some examples, the pin-to-pin range of the mechanical stroke can be between 7.7 inches and 14.8 inches.

[0139] In one or more examples, a damper (e.g., damper 1325, 1425, 1525, or 1625) can be configured to produce a worst-case hinge moment at an end stop of an aircraft (such as end stop 1548 shown in FIG. 15) at a speed less than the maximum allowable impact speed, where the end stop is a component configured to be impacted at a maximum tilt by a tiltable component (such as a swashplate) in each direction of the aircraft, as further described below. In some examples, at a damper piston linear rate of about 12 inches / second, the worst-case hinge moment at the maximum allowable impact speed can be between 2500 lbf and 3000 lbf. Thus, the damper can be configured to provide different amounts of force based on the rate of linear extension or retraction of the actuator to tilt a tiltable swashplate during normal operation.

[0140] Figure 17 An exemplary graph is shown, which shows the damping force of a damper using TPD MD-914-04 damping fluid at various linear extension and retraction rates. In Figure 17 The linear rate of the exemplary damper provided on the X-axis of the graph is the extension or retraction rate of the exemplary damper piston (e.g., the piston of any one of dampers 1325, 1425, 1525, and 1625). A positive value on the X-axis indicates the extension of the damper piston (e.g., damper piston 1318 or 1418), while a negative value on the X-axis indicates the retraction of the damper piston. When configured to tilt a swashplate (e.g., Figure 2When the swashplate 203) tilted actuator (e.g., actuator 1306 or 1406) causes the swashplate to tilt towards the lift configuration (or in the case of actuator shaft breakage, the swashplate tilts towards the lift configuration in an uncontrolled manner), the damper applies a tensile force proportional to the piston speed to the swashplate. Alternatively, when the actuator causes the swashplate to tilt towards the forward flight configuration (or similarly, in the case of actuator shaft breakage, the swashplate tilts towards the lift configuration in an uncontrolled manner), the damper applies a compressive force to the swashplate. Thus, when the swashplate tilts, the damper will apply a force in the direction opposite to the direction of swashplate tilt.

[0141] Figure 17 Includes three different lines 1701-1703, which indicate three different relationships between the linear extension or retraction rate of the damper piston and the damping force. Line 1701 indicates the lower threshold of the damping force at various linear rates, line 1702 indicates the nominal damping force of an exemplary damper at various linear rates, and line 1703 indicates the upper threshold of the damping force at various linear rates. As Figure 17 shown, an exemplary damper can be configured to maintain a damping force of less than 400 lbf at actuator extension and retraction rates of + / - 0.68 inches / second to minimize losses into the actuator. Using Figure 17 the exemplary damper with the MIL-PRF-87257 damping fluid depicted in is configured to handle ultimate loads of 5000 lbf extension and 3000 lbf retraction.

[0142] Selecting a damper that inhibits high-speed tilting of a tiltable swashplate in the case of actuator disconnection can be at least partially based on the characteristics of the swashplate. According to some examples, the tiltable swashplate disclosed herein can have a mass moment of inertia between 5 kg / m² and 10 kg / m². As described below, an exemplary swashplate with a mass moment of inertia of approximately 6.7 kg / m² is used to determine the upper and lower limits of the rotational energy of the swashplate when it impacts the end stop of the aircraft during uncontrolled tilting of the tiltable swashplate.

[0143] According to some examples, the rotational energy of a tiltable swashplate when it impacts the end stop of the aircraft during uncontrolled rotation can be between 0 J and 1000 J. The lower limit of the rotational energy can be between 0 J and 300 J, and the upper limit of the rotational energy can be between 700 J and 1000 J. The lower limit can include the case where the actuator is disconnected and the tiltable swashplate impacts the end stop of the aircraft at a glide speed after minimal acceleration (e.g., acceleration between 0 milliseconds and 100 milliseconds). The upper limit can include the case where the actuator is disconnected and the tiltable swashplate accelerates until it impacts the end stop of the aircraft after disconnection.

[0144] In some examples, the uncontrolled tilting of the tiltable propeller due to the actuator disconnecting can cause the tiltable propeller to impact the end stop of the aircraft at a rate between 500 degrees per second and 1000 degrees per second. According to some examples, the acceptable impact rate of the tiltable propeller hitting the end stop of the aircraft is between 150 degrees per second and 200 degrees per second, which is much lower than the 500 degrees per second to 1000 degrees per second experienced during uncontrolled tilting. In some examples, the allowable impact rate can be less than 100 degrees per second. Thus, the passive damper provided herein can be configured to reduce the impact rate of the tiltable propeller by preventing the propeller from accelerating beyond the acceptable impact rate in the event of an actuator disconnect. Depending on the various requirements of the aircraft, the passive damper can be configured to reduce the impact rate 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. The rate at which the actuator must be able to tilt the tiltable propeller between the lift position and the forward flight position during normal operation can affect the amount by which the damper is configured to reduce the angular rate of the propeller. In some examples, the actuator is configured to tilt the tiltable propeller between the lift position and the forward flight position at a maximum of between 5 degrees per second and 30 degrees per second. In some examples, the actuator is configured to tilt the tiltable propeller between the lift position and the forward flight position at a maximum of 15 degrees per second. In some examples, the actuator is configured to tilt the tiltable propeller at a maximum of 15 degrees per second when hovering (i.e., when stabilizing the tilt angle of the propeller in the lift / hover configuration during normal operation), and at a maximum of 8 degrees per second when translating (i.e., when the propeller is transitioning between the forward flight configuration and the lift / hover configuration during normal operation).

[0145] Results of an exemplary tiltable propeller disconnect simulation analysis are in Figures 18A - 18D shown. It should be understood that Figures 18A - 18D the simulation results depicted in are only intended to provide illustrative performance data of an exemplary passive damping system. The results of the tiltable propeller disconnect simulation analysis should not be construed as restrictive, as those skilled in the art will understand that many different passive dampers can be configured to inhibit the tilt speed of the tiltable propeller described herein without departing from the scope of the claims.

[0146] Figure 18A depicts the external hinge moment and the damping hinge moment acting on the propeller as a function of time starting approximately at the moment of actuator disconnect. The hinge moment provided on the Y-axis of the graph in Figure 18A is the moment acting about the axis of rotation of the propeller, which axis of rotation is about the joint that connects the propeller to the aircraft 100. The axis of rotation of the propeller is shown as axis of rotation 1301 in the example of Figure 13 Figure 18A ​Contains two lines, line 1801 and line 1802. Line 1801 represents the hinge moment of the pitch-changing propeller, and line 1802 shows the hinge moment of the damper. As shown, the hinge moment of the damper can closely track the hinge moment of the pitch-changing propeller. Within 200 milliseconds after the actuator is disconnected, the hinge moments of both the pitch-changing propeller and the damper can increase to between 400 ft-lbf and 600 ft-lbf. As shown in the figure, after the actuator is disconnected, the hinge moment can then briefly decrease and then rise to a maximum value between 800 ft-lbf and 1000 ft-lbf between one second and two seconds.

[0147] Figure 18B A graph depicting the angular velocity of the pitch-changing propeller as a function of time starting approximately at the moment the actuator is disconnected. At Figure 18B The angular rate provided on the Y-axis of the graph is the angular velocity at which the pitch-changing propeller rotates about the axis of rotation of the pitch-changing propeller. The axis of rotation of the pitch-changing propeller is shown as axis 1301 in the example of Figure 13 The angular rate depicted on the Y-axis is the angular rate of the pitch-changing propeller acting under the influence of damping forces after the actuator is disconnected, where the damping forces are provided by passive dampers such as dampers 1325, 1425, 1525, or 1625. As shown in the figure, when the angular rate is controlled by a damper (such as damper 1325, 1425, 1525, or 1625), the angular rate of the pitch-changing propeller after the actuator is disconnected can be configured to remain below an impact rate of 100 degrees per second.

[0148] Figure 18C A graph depicting the linear velocity of the damper piston as a function of time starting approximately at the moment the actuator is disconnected. At Figure 18C The linear velocity provided on the Y-axis of the graph is the linear velocity of an exemplary damper piston traveling along the extension and retraction axis of the damper (e.g., of damper 1325 or 1425). The approximate extension and retraction axis of the damper is shown as axis 1311 and 1411 in Figure 13 and Figure 14 As shown in the figure, shortly after the actuator is disconnected (within 200 milliseconds), the damper piston can reach a linear velocity between 2 inches per second and 3 inches per second. Within the first two seconds after the actuator is disconnected, the linear velocity of the damper piston can fluctuate between 2.0 inches per second and 3.5 inches per second.

[0149] Figure 18D A graph depicting the damping force as a function of time starting approximately at the moment the actuator is disconnected. At Figure 18DThe damping force provided on the Y-axis of the graph is the force that resists the rotation of the swashplate after the actuator is disconnected. As shown, shortly after the actuator is disconnected (within 200 milliseconds), the damping force can reach an initial peak between 2500 lbf and 3500 lbf. Within the first two seconds after the actuator is disconnected, the damping force can fluctuate between 1500 lbf and 5000 lbf.

[0150] Figure 19 and Figure 20 shows a graph of the effective force arm length of an exemplary damper varying with the swashplate tilt angle. Figure 19 The exemplary damper of is installed in a first exemplary configuration, and Figure 20 The exemplary damper of is installed in a second exemplary configuration. At Figure 19 The effective force arm length of the exemplary damper installed in the first configuration provided on the Y-axis of the graph of is the force arm of the damper at various tilt angles between 0 degrees and 100 degrees of the tiltable swashplate. At Figure 20 The effective force arm length of the exemplary damper installed in the second exemplary configuration provided on the Y-axis of the graph of is the force arm of the damper at various tilt angles between 0 degrees and 100 degrees of the tiltable swashplate. As Figure 19 and Figure 20 shown, when the swashplate tilts to different tilt angles, the installation configuration of the damper relative to the swashplate can affect the effective force arm of the damper.

[0151] In one or more examples, any of the systems described above (such as system 202, system 302, system 402, system 502, system 1302, system 1402, and aircraft 100) may include a tilt-rotor locking mechanism, as Figure 21 shown in the exemplary system 2102 of , which shows a side view of the system 2102 in a first configuration according to one or more examples of the present disclosure. The system 2102 can be implemented in an aircraft having a tiltable swashplate, instead of or in addition to a redundant actuator and / or damper mechanism, and can lock the tilt of the swashplate in a proper position.

[0152] The system 2102 is positioned between the boom 2105 and the swashplate 2103 of the aircraft, and may include a tension spring 2145 connected to a pulley 2140 via a connector 2160, where the pulley 2140 is also connected to a pawl 2152 via a connector 2142. The pawl 2152 can selectively engage with a sector gear 2150 based on the movement of the pulley 2140, where the pawl 2152 is configured to move towards the sector gear 2150 when the pulley 2140 moves away from the actuator 2106. The pulley 2140 is connected to the actuator 2106 during normal operation.

[0153] In the case where the actuator 2106 is disconnected from the pulley 2140 (e.g., the actuator 2106 is disconnected from the paddle 2103), the pulley 2140 moves away from the actuator 2106 due to the biasing force from the spring 2145, as Figure 22 shown in the configuration of the system 2102 shown. Although Figure 21 it shows that the pawl 2152 is not engaged with the sector gear 2150 and the pulley 2140 is located near the actuator 2106, Figure 22 it shows that the pawl 2152 is engaged with the sector gear 2150 and the pulley 2140 is no longer located near the actuator 2106. In the case where the actuator 2106 and the pulley 2140 are disconnected, the tension spring 2145 can automatically pull the pulley 2140 away from the actuator 2106, thereby forcing the pawl 2152 to engage with the sector gear 2150. When the pawl 2152 engages with the sector gear 2150, the paddle 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 in the case where the actuator is disconnected from the paddle, without the need for redundant actuators or damping mechanisms. Figure 23 shows a front view of an exemplary system according to one or more examples of the present disclosure Figure 21 .

[0154] In one or more examples, the sector gear 2250 can be a ratchet having ridges that contact the pawl 2252. Such an exemplary configuration is shown in Figure 24A which shows an exemplary ratchet configuration 2202 according to one or more examples of the present disclosure. The ratchet configuration 2202 includes a sector gear 2250 having a plurality of ridges 2253, the ridges having a beveled side and a straight side. The pawl 2252 passes over the beveled side but catches on the straight side. Thus, the paddle will be able to tilt in a first direction (e.g., counterclockwise tilting is allowed in the shown configuration), but not in a second direction (e.g., clockwise tilting is not allowed in the shown configuration). This can be used to allow the paddle to move to a desired failure state tilt angle, such as a lift configuration, where the paddle can still be used during at least a portion of the 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.

[0155] Alternatively, the sector gear and the pawl can be configured to lock the paddle in both directions. Figure 24BAn exemplary locking configuration 2204 is shown having a sector gear 2250 and a locking mechanism 2254, the sector gear having a plurality of ridges 2255 having two straight edges, the locking mechanism engaging the ridges 2255. When the locking mechanism 2254 is forced against the sector gear 2250, the locking mechanism will engage the ridges 2255. Due to the straight edges of the ridges, the locking mechanism 2254 cannot cross the ridges 2255 in either direction, thereby locking the sector gear 2250 (and thus, the propeller) in place.

[0156] Accordingly, systems and methods are described herein for mechanically relating the pitch of a propeller of an aircraft to the blade pitch of the propeller blades. The system enables adjustment of the blade pitch for different operating states of the propeller while avoiding the need for dedicated blade pitch adjustment actuators and the associated cost, weight, and points of failure associated with such dedicated blade adjustment actuators.

[0157] Figure 25 An alternative spring configuration and features for providing "mid-term" inspection capabilities to ensure that the pawl mechanism is not jammed are shown.

[0158] For purposes of explanation, the foregoing description has been made with reference to specific examples. However, the above illustrative discussion is not intended to be exhaustive or to limit the invention to the precise forms disclosed. Given the above teachings, many modifications and variations are possible. The examples were chosen and described in order to best explain the principles of the technology and its practical application. Accordingly, other technicians in the art will be able to best utilize the technology with various modifications that are suitable for the particular purposes contemplated.

[0159] Although the present disclosure and examples have been described fully with reference to the accompanying drawings, it should be noted that various modifications and variations will become apparent to those skilled in the art. Such changes and modifications should be understood to be included within the scope of the present disclosure and examples as defined by the appended claims.

Claims

1. An aircraft, the aircraft comprising: A tiltable propeller that can tilt between a lift position for providing lift to the aircraft and a forward flight position for providing forward propulsion to the aircraft; At least one actuator for adjusting the tilt angle of the tiltable propeller; And At least one passive damper connected to the tiltable propeller and configured to limit the rate of change of the tilt angle of the tiltable propeller.

2. The aircraft according to claim 1, wherein the at least one passive damper comprises a hydraulic cylinder or a pneumatic cylinder.

3. The aircraft according to claim 2, wherein the hydraulic cylinder or the pneumatic cylinder is a balanced non-powered cylinder or an unbalanced cylinder.

4. The aircraft according to any one of the preceding claims, wherein the at least one passive damper is configured to limit the rate of change of the tilt angle of the tiltable propeller in two tilt directions.

5. The aircraft according to any one of the preceding claims, wherein the at least one actuator comprises a linear actuator.

6. The aircraft according to any one of the preceding claims, wherein the at least one actuator comprises a rotary actuator.

7. The aircraft according to any one of the preceding claims, comprising a boom, wherein the tiltable propeller is tiltably mounted to the boom.

8. The aircraft according to claim 7, comprising a fixed rotor mounted to the boom at a fixed position to provide lift.

9. The aircraft according to claim 8, wherein the tiltable propeller is mounted to the front end of the boom, and the fixed rotor is mounted to the rear end of the boom.

10. The aircraft according to any one of claims 7 to 9, wherein the boom comprises a housing, and the at least one passive damper is at least partially received within the housing.

11. The aircraft according to any one of claims 7 to 10, wherein the boom comprises ribs, and the at least one actuator and the at least one passive damper are positioned on opposite sides of the ribs.

12. The aircraft according to any one of claims 7 to 11, wherein the boom is mounted to the wing of the aircraft, inside the end of the wing.

13. The aircraft according to any one of the preceding claims, wherein the tiltable propeller is configured to tilt from the forward flight position to the lift position in an upward direction, and the force vector of the at least one passive damper extends below the tilt axis of the tiltable propeller.

14. The aircraft according to any one of the preceding claims, comprising a plurality of tiltable propellers and a plurality of passive dampers for the plurality of tiltable propellers.

15. The aircraft according to any one of the preceding claims, wherein the tilt angle of the tiltable propeller ranges from at least 90 degrees.

16. The aircraft according to any one of the preceding claims, wherein the aircraft comprises a single actuator for adjusting the tilt angle of the tiltable propeller.

17. An aircraft according to any one of the preceding claims, wherein the at least one passive damper is configured to limit the rate of change of the tilt angle to a predetermined threshold in the case where the tiltable propeller is disconnected from the at least one actuator during forward flight.

18. An aircraft according to any one of the preceding claims, wherein the tiltable propeller is electrically powered.

19. An aircraft according to any one of the preceding claims, wherein the aircraft is a passenger aircraft.

20. A method for controlling an aircraft, the method comprising: receiving, at a controller, a command to adjust the tilt angle of a tiltable propeller, the tiltable propeller being tiltable between a lift position for providing lift to the aircraft and a forward flight position for providing forward propulsion to the aircraft; and controlling at least one actuator according to the command to adjust the tilt angle of the tiltable propeller, wherein at least one passive damper is connected to the tiltable propeller to limit the rate of change of the tilt angle of the tiltable propeller.

21. The method according to claim 20, wherein the method is performed using an aircraft according to any one of claims 1 to 19.

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