Patent Application
20190118942
Not applicable.
Not applicable.
Tiltrotor aircraft are generally operable in a helicopter flight mode to ascend from and/or descend to a landing area and in an airplane flight mode to propel the aircraft forward. The transition from the helicopter flight mode to the airplane flight mode, and vice versa, is generally accomplished by selectively pivoting pivotable rotor assemblies and/or pylons of the aircraft to between a horizontal orientation and a vertical orientation to change the thrust angle of the rotatable aircraft blades. When operated in the airplane mode at high speeds and/or high altitudes, tiltrotor aircraft are subject to various aeroelastic instabilities which may cause damage to the components of the aircraft and/or the aircraft itself and may induce severe vibrations throughout the fuselage of the aircraft. As tiltrotor aircraft continue to evolve and achieve increasing forward flight speeds and/or altitudes, the aeroelastic instabilities experienced by the aircraft may also increase. Thus, in emerging tiltrotor aircraft, controlling aeroelastic instabilities becomes increasingly important to ensure the safety, reliability, and performance of the aircraft.
For a more complete understanding of the present disclosure and the advantages thereof, reference is now made to the following brief description, taken in connection with the accompanying drawings and detailed description.
In this disclosure, reference may be made to the spatial relationships between various components and to the spatial orientation of various aspects of components as the devices are depicted in the attached drawings. However, as will be recognized by those skilled in the art after a complete reading of this disclosure, the devices, members, apparatuses, etc. described herein may be positioned in any desired orientation. Thus, the use of terms such as “above,” “below,” “upper,” “lower,” or other like terms to describe a spatial relationship between various components or to describe the spatial orientation of aspects of such components should be understood to describe a relative relationship between the components or a spatial orientation of aspects of such components, respectively, as the device described herein may be oriented in any desired direction.
Referring now to
In operation, each pivotable rotor assembly 106 operates to rotate the associated rotor system 108 and rotor blades 110 about an axis to generate a thrust to propel the aircraft 100. Additionally, the pivotable rotor assemblies 106 are selectively pivotable between a horizontal orientation and a vertical orientation with respect to the fuselage 102 and wings 104 to adjust the thrust angle and transition the aircraft 100 between the airplane mode and the helicopter mode. Accordingly, the airplane mode is associated with a more horizontally-oriented thrust angle and propelling the aircraft 100 forward in flight, while the helicopter mode is associated with a more vertically-oriented thrust angle and propelling the aircraft 100 to and from a landing area. Therefore, to adjust the thrust angle from more horizontal to more vertical and transition from the airplane mode to the helicopter mode, the pivotable rotor assemblies 106 may be pivoted from the horizontal orientation to the vertical orientation. To adjust the thrust angle from more vertical to more horizontal and transition from the helicopter mode to the airplane mode, the pivotable rotor assemblies 106 may be rotated from the vertical orientation to the horizontal orientation.
Referring now to
In the embodiment shown, the rotatable ring 132 is generally affixed to the rotor mast 120 and rotates with the rotation of the rotor mast 120, while the non-rotatable ring 134 is mounted about the rotor mast 120 and remains stationary with respect to the rotation of the rotor mast 120 and the rotatable ring 132. A plurality of actuators 138 are coupled to the non-rotatable ring 134 of the swashplate assembly 136 and a stationary component and/or portion of an aircraft 100 at an opposing end of each actuator 138. Each of the plurality of actuators 138 is selectively extendable and retractable to control the position, pitch, tilt, angle, orientation, and/or translation of the non-rotatable ring 134, which is then translated to the non-rotatable ring 132 to selectively adjust the pitch of each of the rotor blades 110 of aircraft 100. As the rotatable ring 132 rotates with the rotor mast 120, each pitch link 130 is driven up and down due to the engagement of the rotatable ring 132 with the non-rotatable ring 134. Accordingly, as the rotatable ring 132 rotates, it drives each pitch link 130, which drives each corresponding pitch horn 128 to rotate each rotor grip 126 and associated rotor blades 110. This allows the pitch of each of the rotor blades 110 to be selectively controlled.
Additionally, it will be appreciated that the selective actuation of the actuators 138 is generally controlled by an electrical signal, hydraulic input, mechanical input, and/or a combination of electrical and/or mechanical signals provided by the flight control subsystem 112 within the aircraft 100. Furthermore, as will be discussed later herein, the rotor system 108 comprises at least one sensor and/or gauge configured to provide feedback regarding operational characteristics of aircraft 100 to the flight control subsystem 112. For example, in some embodiments, the rotor system 108 may comprise at least one accelerometer 140 disposed in a shear plane of the rotor hub 122 and configured to detect destabilizing hub shears relating to whirl motion of the rotor system 108. Further, in some embodiments, the rotor system 108 may comprise at least one strain gauge 142 disposed in and/or on the rotor mast 120 also configured to detect the destabilizing hub shears relating to a whirl motion of the rotor system 108 since hub shear causes deflection and/or bending in the rotor mast 120. However, in some embodiments, the rotor system 108 may comprise at least one accelerometer 140 and one strain gauge 142.
Referring now to
Δθ=−tan(delta-3)Δβ.
The pitch/flap coupling caused by the delta-3 angle alters the aerodynamic forces acting on the rotor system 108, which modifies the flapping frequency. The rotor delta-3 angle is used to reduce rotor system 108 flapping amplitudes during gust disturbances or pilot maneuvers. This prevents excessive flapping which can cause high rotor loads and mechanical interferences. However, the delta-3 angle can be adjusted by moving the location of the pitch horn 128 relative to the flapping axis 150. In traditional tiltrotor aircraft, the delta-3 angle is usually set to values near −15 degrees, which provides an adequate level of flapping attenuation. Larger values of delta-3 would reduce flapping even more, but this can aggravate the aeroelastic stability problems described above.
Because the delta-3 coupling alters the flapping frequency of a rotor system 108, it affects the basic rotor system 108 flapping response characteristics, as well as the destabilizing hub shears in the rotor hub 122. This affects both the proprotor aeroelastic instability and the rotor flap-lag instability. For the proprotor aeroelastic stability problem, large negative values of delta-3 angle will increase the magnitude of the destabilizing hub shears. The increase in negative rotor damping will reduce the stability boundary of the aircraft. Likewise, large positive values of delta-3 are beneficial for proprotor aeroelastic stability. Large positive values of delta-3, however, will cause the flapping frequency to increase and approach the rotor in-plane mode frequency. This can lead to a rotor flap/lag instability at high speed. Likewise, large negative values of delta-3 will improve the rotor flap/lag stability by preventing coalescence of these two rotor modes. Thus, a selected design value of delta-3 is a compromise between the requirement for acceptable flapping reduction, good proprotor aeroelastic stability, and acceptable flap/lag stability.
Referring now to
Whirl flutter may be caused by excessive in-plane or “destabilizing” hub shears, which impose a limit on the forward flight speed of an aircraft 100 and are induced by a variety of factors, including aerodynamic forces, high velocity airflow through rotor system 108, gyroscopic forces caused by rotor system 108, mounting stiffness of the rotor blades 110 to the rotor system 108, flapping of the rotor blades 110, in-flight disturbances such as air gusts, and/or the natural flutter frequency of the rotor blades 110 and/or the wings 104 of the aircraft 100. Whirl flutter may occur in both turbo-prop and tiltrotor aircraft 100. However, tiltrotor aircraft 100 are more prone to whirl flutter instability due to rotor blade 110 flapping and/or bending.
As an aircraft 100 operates in the forward-flight airplane mode shown in
Referring now to
The control system 200 comprises at least one accelerometer 204 substantially similar to accelerometer 140 and/or at least one strain gauge 206 substantially similar to strain gauge 142 associated with each rotor system 108 of aircraft 100. The accelerometers 204 and the strain gauges 206 are configured to detect destabilizing hub shears acting on the rotor hub 122 and/or the rotor mast 120. However, in some embodiments, the accelerometers 204 and the strain gauges 206 may be configured to detect other forces indicative of the presence of destabilizing hub shears in the rotor hub 122 and/or the rotor mast 120. Additionally, in some embodiments, the control system 200 may also comprise at least one sensor 208 associated with each wing 104 and configured to detect bending moments in the wing 104 caused by flapping of the wings 104. However, in some embodiments, each wing 104 may comprise a sensor 208 on each of a top and bottom side of the wing 104. Furthermore, it will be appreciated that control system 200 may comprise any combination of accelerometers 204, strain gauges 206, and/or sensors 208.
In operation, when aircraft 100 travels at high forward-flight speeds and/or experiences irregular gusts of wind, the mechanical stability of the rotor system 108 may be exceeded, thereby causing the rotor system 108 to become unstable and experience whirl flutter. The instability of the rotor system 108 may be triggered by destabilizing hub shear acting on the rotor mast 120 and/or the rotor hub 122. Thus, for the control system 200 to provide electronic stability to aircraft 100 by controlling, reducing, and/or eliminating whirl flutter, the control system 200 is configured to detect the destabilizing hub shears. In some embodiments, the control system 200 may employ an accelerometer 204 disposed in the shear plane of each rotor hub 122 of the aircraft 100 to detect the destabilizing hub shears in the rotor hubs 122. The accelerometers 204 measure hub shear in the rotor hubs 122 by detecting acceleration in the shear planes of the rotor hubs 122 caused by vibrational displacement.
In some embodiments, the control system 200 may employ a strain gauge 206 disposed in and/or on each rotor mast 120 of the aircraft 100 to detect the destabilizing hub shears. The strain gauges 206 measure hub shears present in the rotor masts 120 by detecting deflection and/or bending in the rotor mast 120. This may be accomplished since destabilizing hub shears generate deflection and/or bending in the rotor mast 120. Additionally, in some embodiments, the control system 200 may employ a sensor 208 disposed in and/or on each wing 104 of the aircraft to detect bending moments present in the wings 104 of the aircraft 100. However, in some embodiments, each wing 104 may comprise a sensor 208 on each of a top and bottom side of the wing 104. The bending moments in the wings 104 result from flapping of the wings 104 caused by the destabilizing hub shears acting on the rotor hub 122 and/or the rotor mast 120. The phenomenon of wing 104 flapping caused by whirl flutter may be referred to as whirl flap. Thus, the destabilizing hub shears may be determined by the flight control subsystem 202 based on the values of the bending moments detected by the sensors 208. Furthermore, it will be appreciated that control system 200 may employ any combination of accelerometers 204, strain gauges 206, and/or sensors 208 to determine the destabilizing hub shears acting on the rotor mast 120 and/or the rotor hub 122.
The data sensed by the accelerometers 204, strain gauges 206, and/or sensors 208 may be communicated to the flight control subsystem 202. The flight control subsystem 202 is configured to receive data relating to the hub shears from each accelerometer 204, strain gauge 206, and/or sensor 208 in the aircraft 100 and analyze the data to determine operational characteristics that must be adjusted to eliminate the destabilizing hub shears. The flight control subsystem 202 comprises software and/or hardware configured to determine hub shear values from the communicated data and/or analyze the communicated data to determine if the hub shears are potentially harmful to the rotor hub 122, rotor mast 102, other components of the rotor system 108, wings 104, and/or any other component of the aircraft 100. Additionally, the communicated data from the gauges 204, 206 and/or sensors 208 may be further communicated and/or displayed by a display, gauges, and/or warning lights by the flight control subsystem 202 to alert a pilot as to presence of the destabilizing hub shears and/or an action taken by the flight control system 202 in response to the presence of the destabilizing hub shears. Furthermore, alerting the pilot as to the presence of the destabilizing hub shears may allow a pilot to selectively operate the flight control subsystem 202 to further adjust the pitch of the rotor blades 110, adjust the speed of the aircraft 100, and/or take other action to control, reduce, cancel, and/or eliminate the destabilizing hub shears and resulting whirl flutter. However, in some embodiments, the flight control subsystem 202 may automatically adjust the pitch of the rotor blades 110 in response to the detection of the presence of destabilizing hub shears.
The control system 200 generally comprises a fail-safe tiered control system 200 that incorporates a series of responses in response to destabilizing hub shears being detecting by at least one of the gauges 204, 206 and/or sensors 208. More specifically, the flight control subsystem 202 comprises an algorithm that utilizes the communicated data from the gauges 204, 206 and/or sensors 208 to analyze the data and initiate a tiered set of responses when the flight control subsystem 202 detects destabilizing hub shears and/or determines that the hub shears may be harmful to components of the aircraft 100. When a destabilizing hub shear is detected by the gauges 204, 206 and/or sensors 208, the flight control subsystem 202 may first operate the actuators 138 to tilt the swashplate assemblies 136 to adjust the pitch of the rotor blades 110 of each rotor system 108.
Additionally, when tilting the swashplate assemblies 136, the flight control system 202 may adjust the pitch of the rotor blades 110 of each rotor system 108 individually. The rotor blades 110 may be tilted at different angles with respect to other rotor blades 110 of the same rotor system 108 based on the angle the swashplate assembly 136 is tilted. More specifically, the flight control subsystem 202 may detect the hub shears and determine hub shear vibrations that result at a particular frequency and/or range of frequencies that initiate whirl flutter. The flight control subsystem 202 may determine the phase of the hub vibrations and tilt the swashplate assembly 136 in a swirling motion to dampen and/or eliminate the hub shear vibrations to stabilize the rotor systems 108. Accordingly, by detecting the hub shears and determining the characteristics of the hub shear vibrations, control system 200 provides a quicker response than traditional methods of simply sensing beam bending in the wings 104. This is due at least in part to the detected hub shears occurring locally at the rotor hub 122 and/or rotor mast 120 and being detected in real-time as opposed to waiting to detect wing bending caused by increasingly dangerous hub shears.
If attempts to control the whirl flutter created by hub shears is not successful by tilting the swashplate assemblies 136, the flight control subsystem 202 may automatically initiate the next step in the tiered set of responses by reducing engine power and/or torque to the rotor systems 108 to attempt to control whirl flutter. At the same time as reducing the engine power and/or torque, the flight control subsystem 202 may initiate collective braking by greatly increasing the pitch of the rotor blades 110 to a high collective angle (e.g. at least about 10 degrees, 15 degrees, 20 degrees, 25 degrees, 30 degrees, 35 degrees, and/or 45 degrees) to reduce the speed of rotation of the rotor systems 108. In some embodiments, the flight control subsystem 202 may reduce the speed of the rotor systems 108 to a known statically stable speed of rotation of the rotor systems 108. The additional thrust created by the high collective angle and slower rotational speed may induce whirl flutter damping in the rotor systems 108 and reduce and/or eliminate the hub shears and the associated whirl flutter.
Furthermore, the flight control subsystem 202 may utilize the algorithm to determine the stability margins of the rotor systems 108. The flight control subsystem 202 may also determine which parameters have the greatest effect on improving the determined stability margins. This allows for real-time feedback to the flight control subsystem 202, such that the flight control subsystem 202 may comprise feedback-incorporating smart logic that adjusts the tired responses based on this feedback and/or learned stability margins. Still further, the flight control subsystem 202 may also initiate in-flight stability checks by initiating a whirl by tilting the swashplate assemblies 136 and observing responses of the rotor systems 108. Such responses may also be used to adjust the tiered responses of the flight control subsystem when destabilizing hub shears are detected. Furthermore, while control system 200 is discussed in terms of detecting “destabilizing” hub shears that induce whirl flutter, it will be appreciated that control system 200 is configured to detect any level of hub shear present in the rotor hub 122 and/or the rotor mast 120, and flight control subsystem 202 is configured to take appropriate action to control, reduce, and/or eliminate the detected hub shears.
It will be further appreciated that while control system 200 may generally be configured for operation in aircraft 100, control system 200 may comprise substantially similar components and be configured for any proprotor and/or tiltrotor aircraft. For example, control system 200 may be configured for use in aircraft with soft in-plane and/or stiff in-plane rotor hubs 122. Further, control system 200 may be used in aircraft without a swashplate assembly 136, where the flight control subsystem 202 is configured to adjust pitch of the rotor blades 110 using an electric, hydraulic, and/or electro-mechanical actuator and/or any other mechanism in response to detecting destabilizing hub shears acting on a rotor hub 122 and/or rotor mast 120.
Referring now to
At least one embodiment is disclosed, and variations, combinations, and/or modifications of the embodiment(s) and/or features of the embodiment(s) made by a person having ordinary skill in the art are within the scope of this disclosure. Alternative embodiments that result from combining, integrating, and/or omitting features of the embodiment(s) are also within the scope of this disclosure. Where numerical ranges or limitations are expressly stated, such express ranges or limitations should be understood to include iterative ranges or limitations of like magnitude falling within the expressly stated ranges or limitations (e.g., from about 1 to about 10 includes, 2, 3, 4, etc.; greater than 0.10 includes 0.11, 0.12, 0.13, etc.). For example, whenever a numerical range with a lower limit, R1, and an upper limit, Ru, is disclosed, any number falling within the range is specifically disclosed. In particular, the following numbers within the range are specifically disclosed: R=R1+k*(Ru−R1), wherein k is a variable ranging from 1 percent to 100 percent with a 1 percent increment, i.e., k is 1 percent, 2 percent, 3 percent, 4 percent, 5 percent, . . . 50 percent, 51 percent, 52 percent, . . . , 95 percent, 96 percent, 95 percent, 98 percent, 99 percent, or 100 percent. Moreover, any numerical range defined by two R numbers as defined in the above is also specifically disclosed.
Use of the term “optionally” with respect to any element of a claim means that the element is required, or alternatively, the element is not required, both alternatives being within the scope of the claim. Use of broader terms such as comprises, includes, and having should be understood to provide support for narrower terms such as consisting of, consisting essentially of, and comprised substantially of Accordingly, the scope of protection is not limited by the description set out above but is defined by the claims that follow, that scope including all equivalents of the subject matter of the claims. Each and every claim is incorporated as further disclosure into the specification and the claims are embodiment(s) of the present invention. Also, the phrases “at least one of A, B, and C” and “A and/or B and/or C” should each be interpreted to include only A, only B, only C, or any combination of A, B, and C.
