Abstract:
In one embodiment, an apparatus comprises a particle damper for damping a component when the particle damper is attached to the component. The particle damper comprises a plurality of pockets configured to hold a plurality of particles, and the particle damper also comprises an attachment fitting for coupling the particle damper to the component.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This patent application claims the benefit of the filing date of U.S. Provisional Patent Application Ser. No. 62/337,374, filed on May 17, 2016, and entitled “Method and Apparatus for Vibration Attenuation,” the content of which is hereby expressly incorporated by reference. 
     
    
     TECHNICAL FIELD 
       [0002]    This disclosure relates generally to aircraft performance, and more particularly, though not exclusively, to attenuation of aircraft loads and vibrations. 
       BACKGROUND 
       [0003]    Aircraft are subjected to various aerodynamic and operational forces during operation. For example, the aerodynamic forces involved during operation of a rotorcraft may include thrust, drag, lift, and weight. In certain circumstances, aerodynamic and operational forces may increase the structural load on components of an aircraft and may also cause vibration. Excessive loads and vibration during operation of an aircraft (e.g., tail loads and vibration) are undesirable and potentially harmful to the aircraft, as they can negatively impact the structural integrity, mechanical integrity, and performance of the aircraft. For example, loads and vibration can cause components of an aircraft to bend and may reduce the structural integrity and fatigue life of the aircraft. Moreover, vibration is undesirable to passengers of an aircraft, as vibration may cause the aircraft to shake and/or produce loud noise, which may negatively impact the comfort of the passengers. 
       SUMMARY 
       [0004]    According to one aspect of the present disclosure, an apparatus comprises a particle damper for damping a component when the particle damper is attached to the component. The particle damper comprises a plurality of pockets configured to hold a plurality of particles, and the particle damper also comprises an attachment fitting for coupling the particle damper to the component. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0005]      FIGS. 1A-B  and  2  illustrate example aircraft in accordance with certain embodiments. 
           [0006]      FIGS. 3A-E  illustrate an example embodiment of a rotorcraft tail structure with a particle damper. 
           [0007]      FIGS. 4 and 5  illustrate graphs of the performance of an example rotorcraft with and without a tail damper. 
           [0008]      FIG. 6  illustrates a flowchart for an example embodiment of tuning an aircraft particle damper. 
       
    
    
     DETAILED DESCRIPTION 
       [0009]    The following disclosure describes various illustrative embodiments and examples for implementing the features and functionality of the present disclosure. While particular components, arrangements, and/or features are described below in connection with various example embodiments, these are merely examples used to simplify the present disclosure and are not intended to be limiting. It will of course be appreciated that in the development of any actual embodiment, numerous implementation-specific decisions must be made to achieve the developer&#39;s specific goals, including compliance with system, business, and/or legal constraints, which may vary from one implementation to another. Moreover, it will be appreciated that, while such a development effort might be complex and time-consuming, it would nevertheless be a routine undertaking for those of ordinary skill in the art having the benefit of this disclosure. 
         [0010]    In this specification, reference may be made to the spatial relationships between various components and to the spatial orientation of various aspects of components as depicted in the attached drawings. However, as will be recognized by those skilled in the art after a complete reading of the present disclosure, the devices, components, 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 similar 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 components described herein may be oriented in any desired direction. 
         [0011]    Further, the present disclosure may repeat reference numerals and/or letters in the various examples. This repetition is for the purpose of simplicity and clarity and does not in itself dictate a relationship between the various embodiments and/or configurations discussed. 
         [0012]    Example embodiments that may be used to implement the features and functionality of this disclosure will now be described with more particular reference to the attached FIGURES. 
         [0013]      FIGS. 1 and 2  illustrate various example aircraft in accordance with certain embodiments, as discussed further below. 
         [0014]      FIGS. 1A and 1B  illustrate an example embodiment of a rotorcraft  101 .  FIG. 1A  illustrates a side view of rotorcraft  101 , while  FIG. 1B  illustrates a perspective view of rotorcraft  101 . Rotorcraft  101  has a rotor system  103  with a plurality of rotor blades  105 . The pitch of each rotor blade  105  can be managed or adjusted in order to selectively control direction, thrust, and lift of rotorcraft  101 . Rotorcraft  101  further includes a fuselage  107 , anti-torque system  109 , an empennage  111 , and a tail structure  120 . In this example, tail structure  120  can represent a horizontal stabilizer. Torque is supplied to rotor system  103  and anti-torque system  109  using at least one engine. 
         [0015]      FIG. 2  illustrates a perspective view of an example tiltrotor aircraft  201 . Tiltrotor aircraft  201  can include nacelles  203   a  and  203   b , a wing  205 , a fuselage  207 , and a tail structure  220 . In this example, tail structure  220  can represent a vertical stabilizer. Each nacelle  203   a  and  203   b  can include an engine and gearbox for driving rotor systems  211   a  and  211   b , respectively. Nacelles  203   a  and  203   b  are each configured to rotate between a helicopter mode, in which the nacelles  203   a  and  203   b  are approximately vertical, and an airplane mode, in which the nacelles  203   a  and  203   b  are approximately horizontal. 
         [0016]    It should be appreciated that rotorcraft  101  of  FIG. 1  and tiltrotor aircraft  201  of  FIG. 2  are merely illustrative of a variety of aircraft that can be used to implement embodiments of the present disclosure. Other aircraft implementations can include, for example, fixed wing airplanes, hybrid aircraft, unmanned aircraft, gyrocopters, and a variety of helicopter configurations, among other examples. Moreover, it should be appreciated that even though aircraft are particularly well suited to implement embodiments of the present disclosure, the described embodiments can also be implemented using non-aircraft vehicles and devices. 
         [0017]    A rotorcraft (e.g., rotorcraft  101  or rotorcraft  201 ) is subjected to various aerodynamic or operational forces during operation, including thrust, drag, lift, and weight. In certain circumstances, the aerodynamic forces may increase the structural load on components of the rotorcraft and may also cause vibration. In general, vibration may be caused by the rotor(s), engine(s), and/or transmission of the rotorcraft. For example, when the rotor of a rotorcraft is in motion, the structural components of the rotorcraft (e.g., the rotor blades and the tail) are continuously subjected to an oscillating force that may increase the structural load on the rotorcraft components and may lead to vibration. In general, vibration may be low when a rotorcraft is in hover, but may increase during forward flight or transition. For example, during forward flight, vibration may increase with the speed of the rotorcraft, and thus high levels of vibration may result when the rotorcraft is flying at its maximum speed. Moreover, when a rotorcraft is in transition, vibration may increase due to the rotor wake influence on the blade air loads. For example, vibration may increase during descent at low speeds or during thrust at high speeds. In some cases, for example, the tail structure of a rotorcraft may vibrate when the rotorcraft is in flight. For example, in certain flight scenarios, the vortex wake from a rotor may impinge directly on the tail of a rotorcraft and may cause the tail to vibrate vertically. 
         [0018]    Excessive loads and vibration during operation of a rotorcraft (e.g., tail loads and vibration) are undesirable and potentially harmful to the rotorcraft, as they can negatively impact the structural integrity, mechanical integrity, and performance of the rotorcraft. For example, oscillatory loads and vibration can cause the components of a rotorcraft to bend and may reduce the structural integrity and fatigue life of the rotorcraft. Moreover, vibration is undesirable to passengers of a rotorcraft, as vibration may cause the rotorcraft to shake and/or produce loud noise, which may negatively impact the comfort of the passengers. Accordingly, there is a need to control or reduce the loads and vibration of components of rotorcraft and other aircraft during operation. 
         [0019]    In some cases, the loads or vibration of rotorcraft components (e.g., loads and vibration of the tail structure) may be addressed by dynamic tuning of modes, for example, by stiffening structural components or adjusting mass distribution. Stiffening a structural component of a rotorcraft, for example, may be achieved by adding additional composite plies to strengthen the component. Adjusting the mass distribution of a rotorcraft may be achieved by redesigning structural components based on the desired mass distribution, for example, to redistribute the total mass and/or increase the mass of certain components. Alternatively, the mass distribution of a rotorcraft could be adjusted using dedicated tuning masses to increase the mass of certain components. For example, dedicated tuning masses (e.g., depleted uranium) could be added to the tail of a rotorcraft to attenuate oscillatory loads in the tail of the rotorcraft. These approaches, however, can be inefficient and may have various drawbacks. For example, once a rotorcraft has been designed and the manufacturing tooling has been built, redesigning the rotorcraft to stiffen components and/or redistribute mass (e.g., redesigning the tail structure) may require intrusive design and tooling modifications, which may be cost prohibitive and time consuming. Redesigning certain components of the rotorcraft may also trigger a cascading chain of design and tuning modifications to other components of the rotorcraft. Moreover, these approaches can also be ineffective solutions for attenuating the loads and vibration during operation of a rotorcraft. For example, while stiffening a component may increase its strength, stiffening also results in additional weight and may increase the load on the component. Similarly, while increasing the mass of a component may reduce vibration, the increased mass results in additional weight and may increase the load on the component. 
         [0020]    Accordingly, there is a need for an efficient and effective approach of attenuating loads and vibration that occur during operation of rotorcraft and other aircraft. This disclosure describes various embodiments for attenuating aircraft loads and vibration using a particle impact damper (which may also be referred to as a rattle damper). Damping in general is a technique for attenuating the vibrations excited in a particular structure or object. Particle damping may involve the use of freely moving particles in a cavity or enclosure to produce a damping effect. A particle damper, for example, may encapsulate one or more loose particles (e.g., ball bearings) that suppress oscillatory motion by momentum transfer and energy dissipation. 
         [0021]    This disclosure describes various embodiments of using particle dampers to attenuate the loads and vibration that occur during operation of rotorcraft and other aircraft. The embodiments described throughout this disclosure can be used on any aircraft with lightly damped structures, such as aircraft with stabilizers (e.g., horizontal and/or vertical stabilizers), vertical fins, control surfaces, and/or tail skids, among other examples. For example, a particle damper can be incorporated into these lightly damped structures in order to control the oscillatory loads inherent in lightly damped structures excited by the turbulent wake of a rotorcraft. In this manner, a particle damper can be used to reduce the magnitude of the oscillatory load at the bending moment of a particular structure. In some embodiments, for example, a particle damper can be integrated or mounted at the tip or edge of an aircraft structure or internally within the structure (e.g., internally within a spar). For example, in some embodiments, a particle damper can be integrated internally or externally on a rotorcraft stabilizer (e.g., a horizontal stabilizer or tail), fin, or control surface, in order to attenuate internal structural loads, loads at structural interfaces, and vibration. For example, particle dampers can be incorporated in the tail of a rotorcraft (e.g., a particle damper at each end of the tail spar) to control the oscillatory loads and vibration in the tail without meaningfully increasing the weight of the rotorcraft, thus enabling the design of lighter tail structures. Furthermore, because a particle damper can be implemented internally within a structure and/or at the tip or edge of the structure, a particle damper can be used to attenuate loads and vibration for an aircraft without disturbing the aerodynamics and airflow of the aircraft. Moreover, in some embodiments, a particle damper can include features that enable tuning (e.g., multiple pockets, adjustable walls, different particle sizes, and so forth) in order to optimize load and vibration attenuation for minimal weight. 
         [0022]    The embodiments described throughout this disclosure provide numerous technical advantages, including using particle dampers to control, reduce, and/or attenuate loads and vibration during operation of a rotorcraft (e.g., tail loads and vibration), which may improve its safety, reliability (e.g., reliability of avionics and mechanical equipment), and fatigue life (e.g., the lifespan of airframe structural components), and may also improve passenger comfort. The described embodiments are also lightweight, flexible (e.g., they can be tuned and adjusted for different aircraft), and can be implemented without disturbing the aerodynamics and airflow of the aircraft. Moreover, the described embodiments can be implemented on an aircraft even after it has been designed and the manufacturing tooling has been built, thus avoiding intrusive design and tooling modifications that are cost prohibitive and time consuming. 
         [0023]    Example embodiments for attenuating loads and vibration of rotorcraft and other aircraft are described below with more particular reference to the remaining FIGURES. 
         [0024]      FIGS. 3A-E  illustrate an example embodiment of a rotorcraft tail structure  300  with a particle damper  330 . Tail structure  300  or a variation thereof, for example, could be used as the tail of a rotorcraft (e.g., the tail of rotorcraft  101  or rotorcraft  201  from  FIGS. 1 and 2 ). In some embodiments, for example, tail structure  300  could be used as a horizontal stabilizer for a rotorcraft. Tail structure  300  may include an elevator for control and stability of a rotorcraft. Moreover, in some embodiments, particle dampers  330  may be incorporated in tail structure  300  to attenuate loads and vibration in tail structure  300  during operation of the rotorcraft, as described further below. 
         [0025]      FIG. 3A  illustrates the right-side portion of a rotorcraft tail structure  300 . In various embodiments, however, tail structure  300  may be symmetrical, and thus the left-side portion of tail structure  300  may be similar to the right-side portion illustrated in  FIG. 3A . 
         [0026]    In the illustrated embodiment, tail structure  300  includes leading edge  303 , trailing edge  305 , inboard end  307 , and outboard end  309 . Leading edge  303  may be an edge of tail structure  300  that faces towards the front of a rotorcraft, while trialing edge  305  may be an edge of tail structure  300  that faces towards the back of the rotorcraft. Inboard end  307  may be a portion of tail structure  300  near the inner or middle portion of the tail structure  300 , while outboard end  309  may be a portion of tail structure  300  near the outer edge of tail structure  300 . Tail structure  300  further includes a tip cap  332  and light  333 . Tip cap  332  may be a sacrificial and/or removable tip cap used for protecting the outboard end  309  of tail structure  300 . Light  333  may be a light source used for visibility and/or navigation purposes. 
         [0027]    In the illustrated embodiment, tail structure  300  also includes a particle damper  330  at the outboard end  309  of the tail spar. Moreover, while  FIGS. 3A-E  only illustrate the right-side portion of tail structure  300 , a particle damper  330  may be similarly included at the outboard end of the tail spar on the left-side portion of tail structure  300 . Particle damper(s)  330  may be used to attenuate loads and vibration in tail structure  300  during operation of a rotorcraft. For example, during operation, a rotorcraft is subjected to various aerodynamic and operational forces that may increase the structural load and/or cause vibration in certain components of the rotorcraft, such as the tail structure  300 . When a rotorcraft is in flight, for example, its tail structure  300  may be continuously subjected to an oscillating force and other operational forces, such as forces created by the rotor(s), engine(s), and/or transmission of the rotorcraft. For example, in certain flight scenarios, the vortex wake from a rotor may impinge directly on the tail structure  300  of a rotorcraft. These various forces may increase the structural load and/or cause vibration in tail structure  300 . For example, in some cases, vibration in tail structure  300  may be low when a rotorcraft is in hover, but may increase during forward flight or transition. Excessive loads and vibration in tail structure  300  are undesirable and potentially harmful to the rotorcraft, as they can negatively impact the structural integrity, mechanical integrity, and performance of the rotorcraft. For example, oscillatory loads and vibration can cause tail structure  300  and/or other components of a rotorcraft to bend and may reduce the structural integrity and fatigue life of the rotorcraft. Moreover, vibration is undesirable to passengers of a rotorcraft, as vibration may cause the rotorcraft to shake and/or produce loud noise, which may negatively impact the comfort of the passengers. Accordingly, in some embodiments, particle damper  330  may be incorporated in tail structure  300  to attenuate the loads and vibration in tail structure  300  during operation of a rotorcraft. Particle damper  330  may involve the use of freely moving particles in a cavity or enclosure to produce a damping effect. For example, particle damper  330  may encapsulate one or more loose particles (e.g., ball bearings) that suppress oscillatory motion by momentum transfer and energy dissipation. Moreover, in the illustrated embodiment, particle damper  330  includes a removable cover  331  that can be removed in order to add or remove particles to or from particle damper  330 , for example, for tuning and adjustment purposes. 
         [0028]      FIGS. 3B and 3C  illustrate a portion of tail structure  300  of  FIG. 3A , but with the cover  331  of particle damper  330  removed. In the illustrated examples, particle damper  330  includes a plurality of pockets  340   a - c .  FIG. 3B  illustrates the pockets  340   a - c  when they are empty, while  FIG. 3C  illustrates the pockets  340   a - c  after they have been filled with a plurality of particles  350   a - c . In some embodiments, for example, a pocket  340  may be a chamber or cavity that is located on an outboard end  309  of tail structure  300 . For example, in some embodiments, one or more pockets  340  could be located in tip cap  332  of tail structure  300 . Moreover, a pocket  340  can contain a plurality of particles  350 . In some embodiments, particles  350  may be generally spherical in shape and can be manufactured out of a ceramic or a metal such as steel or tungsten. For example, in some embodiments, particles  350  may be or may include a plurality of ball bearings. 
         [0029]    As explained above, during operation of a rotorcraft, various aerodynamic and operational forces may increase the structural load and/or cause vibration in tail structure  300 . When particles  350  in pockets  340  are excited by vibrations during operation of the rotorcraft, the movement of particles  350  within pockets  340  causes damping of the vibration in tail structure  300 . For example, the impact of particles  350  on each other and on the walls of pockets  340  (e.g., the top cover  331 , sides, and bottom of pockets  340 ), the friction between each particle  350 , and the friction between particles  350  and the walls of pockets  340  cause energy dissipation, which reduces the amplitude of the vibration of tail structure  300 . 
         [0030]    Moreover, in some embodiments, a particle damper  330  may be designed as a standalone component that can be fastened onto a particular structure or component. In some embodiments, for example, a standalone particle damper  330  may be attached to a particular structure or component using one or more attachment mechanisms or fittings, such as mechanical fasteners and fittings (e.g., threaded holes, plugs, anchors, threaded fasteners, screws, bolts, nuts, washers) and/or any other suitable attachment mechanisms. For example, in some embodiments, a particle damper  330  could be designed as a standalone component that can be fastened to an outboard end  309  of a tail structure  300  of a rotorcraft. Thus, in some embodiments, a particle damper  330  may extend the length of a tail spar (e.g., by 4 inches in some embodiments), which in turn may generate more lift during operation of a rotorcraft. In some embodiments, a particle damper  330  could also be embedded in a tip cap  332  that can be attached to an outboard end  309  of a tail structure  300 . Moreover, in some embodiments, a particle damper  330  may span the full width or chord of a tail spar, while in other embodiments a particle damper  330  may span less than the full width or chord of the tail spar (e.g., the particle damper  330  may be truncated such that it only spans a percentage (e.g., 70%) of the full width or chord of the tail spar). In addition, in some embodiments, a particle damper  330  could be designed with removable pocket plugs  340  to facilitate tuning and adjustment. Moreover, particle dampers  330  can be designed using varying numbers of pockets  340 , and the pockets  340  can be designed using varying sizes, dimensions, shapes (e.g., circular or square), and/or locations. In addition, the pockets  340  can each be filled with particles  350  of varying sizes and shapes. In some cases, for example, a pocket  340  could be filled with size #F steel shot ball bearings (e.g., ball bearings with a diameter of approximately 0.22 inches or 5.6 millimeters), while in other cases, a pocket  340  could be filled with size #2 steel shot ball bearings (e.g., ball bearings with a diameter of approximately 0.150 inches or 3.8 millimeters). 
         [0031]    These various embodiments allow a particle damper  330  to be designed for a particular aircraft and/or aircraft component (e.g., a rotorcraft tail structure  300 ) after the aircraft and/or aircraft component have already been designed and manufactured, thus enabling the design of the particle damper  330  to be tailored for its particular use. Similarly, these various embodiments also facilitate tuning and adjustment of a particle damper  330  for the particular aircraft or component that the particle damper  330  is used with. A particle damper  330  with multiple pockets  340 , for example, may facilitate tuning and adjustment by allowing the particles  350  to be spread across the various pockets  340  of the particle damper  330 . For example, if more velocity from the particles  350  is needed in order to counteract the loads and vibration in the tail structure  300  of a rotorcraft during flight, the particles  350  may be spread across the pockets  340  of the particle damper  330  to provide more room for movement within each pocket  340 , which may increase the velocity and impact of the particles  350 . Moreover, each pocket  340  of a particle damper  330  can be filled with varying number of particles  350 , and the particles  350  in each pocket  340  can be varying sizes and/or shapes. 
         [0032]    In this manner, the performance or effectiveness of a particle damper  330  can be ascertained based on testing and simulation (e.g., computer-based simulations, hangar testing, and flight testing), and the design and/or configuration of the particle damper  330  can be subsequently adjusted or tuned, as appropriate. For example, in some embodiments, a particle damper  330  can be adjusted or tuned by removing its cover  331  and adjusting the number, size, and/or shape of particles  350  within each pocket  340  of the particle damper  330 . A particle damper  330  with removable pocket plugs  340  can be adjusted in a similar manner by removing the pocket plugs  340 , adjusting the particles  350  within each pocket plug  340 , and inserting the pocket plugs  340  back into the particle damper  330 . 
         [0033]    In other embodiments, a particle damper  330  may be permanently embedded into the design of a particular structure or component. For example, in some embodiments, the tail structure  300  of a rotorcraft could be designed with particle dampers  330  embedded at each outboard end  309  of the tail structure  300 . In these embodiments, a particle damper  330  may still be designed with features that facilitate adjustment and/or tuning, such as removable pocket covers (or removable pockets) for adjusting the particles  350  within each pocket  340 , pockets  340  with adjustable sizes and/or walls, and so forth. 
         [0034]      FIGS. 3D and 3E  illustrate a portion of tail structure  300  of  FIG. 3A  without tip cap  332 . Moreover, in  FIG. 3D , light  333  is included in tail structure  300 , while in  FIG. 3E , light  333  has been removed from tail structure  300 . In some embodiments, for example, a plug (not shown) filled with a plurality of particles  350  can be inserted into the hollow portion of the spar  339  of tail structure  300 . 
         [0035]      FIGS. 4 and 5  illustrate graphs of the performance of an example rotorcraft with and without a tail damper. 
         [0036]      FIG. 4  illustrates a graph  400  plotting the acceleration  420  (e.g., response attenuation) and frequency  410  of the tail of a rotorcraft during a hangar test. A hangar test, for example, may involve shaking a rotorcraft in a test hangar and measuring the acceleration  420  of the tail of the rotorcraft at varying frequencies  410 . In graph  400 , the acceleration  420  is represented as units of gravity (G) and the frequency  410  is represented as hertz (Hz). 
         [0037]    Plot  401  corresponds to an example rotorcraft without a tail damper. As illustrated by plot  401 , the acceleration  420  of the tail reaches a peak of approximately 7 G at a frequency of approximately 15.5 Hz. This data indicates that turbulence coming off the front of the rotorcraft causes the tail to respond or vibrate with high acceleration (e.g., approximately 7 G) at the natural frequency of the tail (e.g., approximately 15.5 Hz). This high peak acceleration in the tail can cause harmful structural loads in the rotorcraft airframe. Moreover, stiffening the tail and/or adding mass may be ineffective solutions for reducing the structural loads, as those approaches may simply adjust the natural frequency  410  of the tail without decreasing the amplitude or peak of the acceleration  420  at that frequency. For example, stiffening the tail may increase its natural frequency  410 , while adding mass to the tail may decrease its natural frequency  410 , but neither approach decreases the amplitude or peak of the acceleration  420  at the natural frequency. Thus, by stiffening or adding mass, a similar peak acceleration  420  would still occur in the tail but at a different frequency  410 . Moreover, both approaches increase the weight of the tail and thus also increase inertia, which can compound the problem of structural loads in the airframe of the rotorcraft. 
         [0038]    Plot  402  corresponds to an example rotorcraft with a tail damper. Damping is an approach that can be used to reduce the peak acceleration  420  without significantly increasing the weight of the tail of the rotorcraft. As illustrated by plot  402 , using a tail damper decreases the peak acceleration  420  of the tail from approximately 7 G to less than 3 G, which is a reduction of approximately 60%. This data indicates that the tail vibration at the natural frequency of the tail (e.g., as caused by turbulence) is counteracted by the moving particles (e.g., ball bearings) in the tail damper. Accordingly, damping the tail vibration in this manner reduces the structural loads in the rotorcraft airframe. 
         [0039]      FIG. 5  illustrates a graph  500  of the load attenuation in the tail versus the airspeed of a rotorcraft during a test flight. Graph  500  plots the oscillatory bending moment  520  (e.g., load attenuation) of the tail of the rotorcraft at varying speeds  510 . In graph  500 , the speed  510  of the rotorcraft is represented using the Knots True Airspeed (KTAS) and the oscillatory bending moment  520  is represented using inches-pounds (in-lbs). A bending moment is a reaction induced in a structural component when an external force is applied to the structural component and causes it to bend. If the force applied to a particular structural component (e.g., a rotorcraft tail structure) exceeds the structural limit of the component (e.g., the external force exceeds the force that can tolerated by the structural component), the structural integrity of the component may be weakened or damaged, thus reducing the fatigue life of the component. In graph  500 , the oscillatory bending moment  520  represents the bending moment at the root of the tail of a rotorcraft when the tail is shaking vertically during flight. Accordingly, in order to preserve the structural integrity of the tail of a rotorcraft, it is beneficial to minimize the oscillatory bending moment  520  of the tail during flight. 
         [0040]    Plot  501  corresponds to an example rotorcraft without a tail damper, while plot  502  corresponds to an example rotorcraft with a tail damper. As illustrated by plots  501  and  502 , the oscillatory bending moment  520  in the tail increases as the speed  510  of the rotorcraft increases, which shows that the vibration and loads in the tail of the rotorcraft increase as the rotorcraft flies faster. However, the oscillatory bending moment  520  in the tail is lower with a tail damper (plot  502 ) than it is without a tail damper (plot  501 ). Thus, this data demonstrates that the oscillatory load in the tail is reduced by the tail damper. Moreover, the difference in the oscillatory bending moment  520  with and without a tail damper generally increases as the speed  510  of the rotorcraft increases. Thus, this data demonstrates that a tail damper provides the most significant benefits at higher speeds. For example, the difference in the oscillatory bending moment  520  with and without a tail damper is larger at higher speeds than at lower speeds. Accordingly, a tail damper provides a more significant reduction in loads and vibration at higher speeds (e.g., when the loads and vibration are significant) than at lower speeds (e.g., when the loads and vibration are relatively minimal or steady). 
         [0041]      FIG. 6  illustrates a flowchart  600  for an example embodiment of tuning an aircraft particle damper (e.g., a particle damper on the tail of a rotorcraft). Flowchart  600  may be implemented, in some embodiments, using the particle damper embodiments described throughout this disclosure (e.g., particle damper  330  of  FIGS. 3A-E ). 
         [0042]    The flowchart may begin at block  602  by configuring a particle damper using an initial configuration. In some embodiments, for example, a particle damper may be designed with features that facilitate tuning and/or adjustment, such as removable pocket covers (or removable pockets) for adjusting the particles within each pocket, pockets with adjustable sizes and/or walls, and so forth. For example, configuring a particle damper may involve adjusting the number, size, and/or shape of particles within each pocket of the particle damper. In some cases, the initial configuration of a particle damper may be identified based on computer-based simulations and testing, prior configurations, and/or a baseline initial configuration. 
         [0043]    The flowchart may then proceed to block  604  to perform testing on the particle damper. For example, the performance or effectiveness of the particle damper (e.g., its effectiveness for reducing loads and vibration) can be ascertained from tests and simulations, such as computer-based simulations, hangar testing, and flight testing. In some embodiments, for example, computer-based simulations may be used to identify the initial configuration of the particle damper, while hangar testing and/or flight testing may be used to further tune, adjust, and/or validate the configuration of the particle damper. For example, after identifying an initial configuration (e.g., based on computer-based simulations), hangar testing may then be iteratively performed to continue tuning the configuration of the particle damper. In some cases, once a suitable configuration for the particle damper has been identified from the hangar testing, flight testing may then be used to validate and/or adjust the identified configuration. 
         [0044]    The flowchart may then proceed to block  606  to determine whether a suitable configuration for the particle damper has been identified. For example, in some embodiments, the testing performed at block  604  may reveal whether, and to what extent, the particle damper reduces or attenuates the loads and vibration in the aircraft (or in the aircraft tail or other structure). If the testing performed at block  604  reveals that the particle damper reduces or attenuates the loads and vibration by a desired threshold, then at block  606  it is determined that the testing was successful and thus a suitable configuration for the particle damper has been identified. At this point, the flowchart may be complete. However, if the testing performed at block  604  reveals that the particle damper fails to reduce or attenuate the loads and vibration by a desired threshold, then at block  606  it is determined that the testing was unsuccessful and thus no suitable configuration for the particle damper has been identified. The flowchart may then proceed to block  608  to tune or adjust the configuration of the particle damper, as described below. 
         [0045]    At block  608 , the configuration of the particle damper is tuned or adjusted. For example, in some embodiments, the configuration of the particle damper may be tuned or adjusted based on the testing performed at block  604 . For example, if the testing from block  604  reveals that more velocity from the particles is needed in order to counteract the loads and vibration in a rotorcraft tail structure, the particles may be spread across the pockets of the particle damper to provide more room for movement within each pocket, which may increase the velocity and impact of the particles. Alternatively, if the testing reveals that more particles are needed to counteract the loads and vibration in the tail structure, then more particles may be added to one or more pockets of the particle damper. 
         [0046]    After the configuration of the particle damper has been tuned and/or adjusted at block  608 , the flowchart may then proceed back to block  604  to continue testing, adjusting, and/or tuning the particle damper until a suitable or optimal configuration is identified. In this manner, the performance or effectiveness of a particle damper can be ascertained based on testing and simulation (e.g., computer-based simulations, hangar testing, and flight testing), and the design and/or configuration of the particle damper can be subsequently adjusted or tuned, as appropriate, until a suitable or optimal configuration is identified. 
         [0047]    At this point, the flowchart may be complete. In some embodiments, however, the flowchart may restart and/or certain blocks may be repeated. For example, in some embodiments, the flowchart may proceed to step  608  to continue adjusting and tuning the configuration of the particle damper for optimal performance. Alternatively, in other embodiments, the flowchart may restart at step  602  to restart the configuration process for the particle damper in connection with another aircraft or aircraft component that the particle damper may be used with. 
         [0048]    The flowcharts and diagrams in the FIGURES illustrate the architecture, functionality, and operation of possible implementations of various embodiments of the present disclosure. It should also be noted that, in some alternative implementations, the function(s) associated with a particular block may occur out of the order specified in the FIGURES. For example, two blocks shown in succession may, in fact, be executed substantially concurrently, or the blocks may sometimes be executed in the reverse order or alternative orders, depending upon the functionality involved. 
         [0049]    Although several embodiments have been illustrated and described in detail, numerous other changes, substitutions, variations, alterations, and/or modifications are possible without departing from the spirit and scope of the present invention, as defined by the appended claims. The particular embodiments described herein are illustrative only, and may be modified and practiced in different but equivalent manners, as would be apparent to those of ordinary skill in the art having the benefit of the teachings herein. Those of ordinary skill in the art would appreciate that the present disclosure may be readily used as a basis for designing or modifying other embodiments for carrying out the same purposes and/or achieving the same advantages of the embodiments introduced herein. For example, certain embodiments may be implemented using more, less, and/or other components than those described herein. Moreover, in certain embodiments, some components may be implemented separately, consolidated into one or more integrated components, and/or omitted. Similarly, methods associated with certain embodiments may be implemented using more, less, and/or other steps than those described herein, and their steps may be performed in any suitable order. 
         [0050]    Numerous other changes, substitutions, variations, alterations, and modifications may be ascertained to one of ordinary skill in the art and it is intended that the present disclosure encompass all such changes, substitutions, variations, alterations, and modifications as falling within the scope of the appended claims. 
         [0051]    In order to assist the United States Patent and Trademark Office (USPTO), and any readers of any patent issued on this application, in interpreting the claims appended hereto, it is noted that: (a) Applicant does not intend any of the appended claims to invoke paragraph (f) of 35 U.S.C. §112, as it exists on the date of the filing hereof, unless the words “means for” or “steps for” are explicitly used in the particular claims; and (b) Applicant does not intend, by any statement in the specification, to limit this disclosure in any way that is not otherwise expressly reflected in the appended claims.