Abstract:
An actuator is disclosed for an active vibration and noise control system in an aircraft. The actuator is configured to attach a vibrating component, such as a gearbox mounting foot, to an aircraft support structure. The actuator is actuated by the control system to control the transmission of vibratory loads from the gearbox foot. The actuator includes a housing and a mounting member. The housing mounts to the aircraft. The mounting member attaches to the vibrating component within the aircraft. A piston arrangement is attached to the housing and includes a sleeve located within the housing, and a piston slidably disposed within the sleeve. A bearing assembly engages the mounting member with the piston. The bearing assembly includes a first bearing located between the mounting member and an inner bearing member. The first bearing is adapted to transmit axial motions between the mounting member and the piston. A second bearing is located between the inner bearing member and the housing and is adapted to transmit moment and shear loads from the mounting member to the housing. A third bearing is located between the inner bearing member and the piston, and is adapted to permit rotational movement of the inner bearing member relative to the piston.

Description:
RELATED APPLICATION  
       [0001]    This application is related to provisional application entitled “Actuator for an Active Transmission Mount Isolation System”, filed Sep. 15, 2000, Ser. No. 60/233,213. 
     
    
       [0002] The Government has rights to the invention pursuant to government contract N00019-96-C-2079 awarded by the United States Naval Research Laboratory. 
     
    
     
       FIELD OF THE INVENTION  
         [0003]    The present invention relates to an actuator for an aircraft and, more particularly, to an actively controlled actuator for reducing vibratory transmissions from a gearbox mount to a support structure.  
         BACKGROUND OF THE INVENTION  
         [0004]    Helicopter main rotor lift and rotor driving torque produce reaction forces and moments on the helicopter main gearbox. The aircraft structure which supports the gearbox, e.g., transmission beams, are designed to react these loads and safely and efficiently transmit these primary flight loads to the airframe.  
           [0005]    In addition to the primary flight loads, the aircraft is also subjected to vibratory loads originating from the main rotor system and acoustic loads generated by clashing of the main transmission gears. These vibratory and acoustic loads produce vibrations and noise within the aircraft that cause discomfort to the passengers and crew. Low frequency rotor vibrations are a leading cause of maintenance problems in helicopters. Furthermore, as the aircraft reaches its maximum forward speed in level flight, the vibratory loads become very large, thus, producing increasingly high vibrations.  
           [0006]    Many attempts have been made over the years to alleviate or reduce these vibratory loads and the resulting vibration and audible noise that develops within the aircraft cabin. A considerable amount of those attempts have been directed toward passive control of the vibrations. Some of the passive solutions have involved changes in rotor blade design to reduce the blade response to the periodic loading it experiences in forward flight. Other passive attempts have been directed toward reducing the transmission of vibratory and acoustic noise into the airframe or from the airframe into the cabin. For example, absorbing blankets have been incorporated between the airframe and the cabin interior for attenuating acoustic energy before it enters the cabin section. Another passive attempt involves the installation of low frequency vibration absorbers around the aircraft that are tuned to the vibration frequency of interest. The tuning is typically at a frequency of NP where N is the number of blades and P is the rotor rotational speed in cycles per second. Tuned absorbers have also been incorporated onto the main transmission support beams to produce a vibration impedance mismatch on and/or near the foot of the transmission.  
           [0007]    One example of a passive vibration absorber is disclosed in U.S. Pat. No. 4,362,281 which relates to a pylon mounting system for supporting a helicopter gearbox. The pylon support is attached to the airframe substructure through resilient couplings or mounts. The couplings include elastomeric bushings which provide a soft resilient attachment between the pylon support and the airframe.  
           [0008]    The above described passive solutions to reduce noise and vibration transmission have generally proven to be heavy and, consequently, not structurally efficient. These prior attempts also allow excessive motion of the gearbox causing gearbox-to-engine shaft misalignment under quasi-steady flight loads exerted by the main rotor.  
           [0009]    There has been some recent attempts at producing active vibration and noise control systems. These systems monitor the status of the aircraft and/or the vibration producing component and attempt to command countermeasures to reduce the noise and vibrations. Active vibration and noise control systems are considered to be better at reducing aircraft vibrations and noise since the systems can be designed to counteract or cancel the vibratory and acoustic loads at or near the structural interface between the transmission and airframe, thus, preventing undesirable loads from entering the airframe.  
           [0010]    U.S. Pat. Nos. 4,819,182 and 5,219,143 disclose one attempt at providing an active vibration control system. This system includes a plurality of vibration sensors, e.g., accelerometers, that are located at strategic places throughout the aircraft and provide signals to an adaptive control unit. The control unit provides signals to electro-hydraulic actuators that are located within a series of struts which support the gearbox. The actuators produce controlled forces which attempt to minimize vibration at the sensed locations.  
           [0011]    Another active control system is discussed in U.S. Pat. No. 5,588,800. This active control system is mounted within a helicopter rotor blade and includes actuatable flaps on the rotor that are controlled to reduce the blade vortex interaction and/or vibratory loads transmitted to the airframe.  
           [0012]    Many of the active control systems that are currently being evaluated or proposed utilize hydraulically operated actuators to provide the counteracting forces for damping the sensed vibratory loads. These actuators include a piston arrangement that is attached to a mounting stub through a ball or universal joint. These types of joints, however, tend to bind under high load, especially high vibratory loads. Also, these actuators incorporate conventional internal seals which are not suitable for vibration and noise. As such, the seals quickly wear out and are not very efficient at attenuating acoustic noise.  
           [0013]    Recently, a vibration reduction system was incorporated into an EH-101 aircraft manufactured by Westland Helicopters. The system included a semi-stiff strut mounted in parallel with the piston load to carry the large quasi-steady flight loads. In this design, however, the strut provided a path for undesirable high frequency acoustic vibrations.  
           [0014]    A need therefore exists for an improved actuator for use in an active vibration control system to minimize NP vibratory and high frequency acoustic transmissions from a vibrating component into the aircraft airframe.  
         SUMMARY OF THE INVENTION  
         [0015]    The present invention relates to an active vibration and noise control system for controlling the transmission of vibratory loads from a vibrating component. The control system includes an actuator that is designed to attach the vibrating component, such as a gearbox mounting foot, to a support structure. The actuator is actuated by the control system to control the transmission of steady-state and transient loads from the vibrating component.  
           [0016]    In one embodiment of the invention, the actuator includes a housing and a mounting member. The housing mounts to the aircraft. The mounting member attaches to the vibrating component within the aircraft. A lap-fit piston arrangement is attached to the housing and includes a sleeve located within the housing, and a piston slidably disposed within the sleeve.  
           [0017]    An elastomeric bearing assembly engages the mounting member with the piston. The bearing assembly includes a first bearing located between the mounting member and an inner bearing member. The first bearing is adapted to transmit axial loads between the mounting member and the piston. A second bearing is located between the inner bearing member and the housing and is adapted to transmit moment and shear loads from the mounting member to the housing. A third bearing is located between the inner bearing member and the piston, and is adapted to permit rotational movement of the inner bearing member relative to the piston.  
           [0018]    A diaphragm separates the piston from the bearing assembly so as to inhibit hydraulic fluid from contacting the bearing assembly.  
           [0019]    The control system includes a processor which receives a plurality of signals representing the vibratory state of the component or airframe and the position state of the actuators. The processor sends signals to a hydraulic actuation system for controlling the movement of the actuators.  
           [0020]    The foregoing and other features and advantages of the present invention will become more apparent in light of the following detailed description of the preferred embodiments thereof, as illustrated in the accompanying figures. As will be realized, the invention is capable of modifications in various respects, all without departing from the invention. Accordingly, the drawings and the description are to be regarded as illustrative in nature, and not as restrictive. 
       
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0021]    For the purpose of illustrating the invention, the drawings show a form of the invention which is presently preferred. However, it should be understood that this invention is not limited to the precise arrangements and instrumentalities shown in the drawings.  
         [0022]    [0022]FIG. 1 is a schematic representation of a transmission arrangement incorporating the present invention in a helicopter.  
         [0023]    [0023]FIG. 2 is a schematic representation of one embodiment of a control system according to the present invention for controlling a plurality of actuators adjacent to a helicopter gearbox.  
         [0024]    [0024]FIG. 3 is an isometric view of one embodiment of an actuator made in accordance with the present invention.  
         [0025]    [0025]FIG. 4 is a cross-sectional view of the actuator taken along lines  4 - 4  in FIG. 3. 
     
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS  
       [0026]    While the invention will be described in connection with one or more preferred embodiments, it will be understood that it is not intended to limit the invention to those embodiments. On the contrary, it is intended that the invention cover all alternatives, modifications and equivalents as may be included within its spirit and scope as defined by the appended claims.  
         [0027]    Certain terminology is used herein for convenience only and is not to be taken as a limitation on the invention. For example, words such as “upper,” “lower,” “left,” “right,” “horizontal,” “vertical,” “upward,” and “downward” merely describe the configuration shown in the figures. Indeed, the components may be oriented in any direction and the terminology, therefore, should be understood as encompassing such variations unless specified otherwise.  
         [0028]    Referring now to the drawings, wherein like reference numerals illustrate corresponding or similar elements throughout the several views, FIG. 1 illustrates a transmission arrangement  10  for a helicopter. The transmission arrangement  10  includes a gearbox  12  which is connected to a helicopter rotor head (not shown). The gearbox  12  is supported by an airframe (generically shown and identified by the reference numeral  14 ). The gearbox is connected to drive train  16 . The gearbox  12  includes a plurality of mounting feet  18  which are typically attached directly to the aircraft structure  14 . However, in the present invention, an active transmission mount  20  is located between at least one mounting foot  18  and the airframe structure  14 . More preferably, there is an Active Transmission Mount (ATM)  20  located between each mounting foot  18  and the airframe structure  14  (i.e, above or below the mounting foot  18  depending on the mounting configuration). There is also preferably an ATM  20  mounted on the lateral side of each transmission mounting foot  18  (FIG. 2) and attached to the airframe  14 .  
         [0029]    Referring now to FIG. 2, one embodiment of the present invention is schematically illustrated for controlling the transmission of vibration from the mounting feet  18  of the gearbox  12  to the airframe  14  (not shown in FIG. 2). In this embodiment, which is a view looking down on the gearbox  12 , the ATMs  20  are shown positioned on the side of and attached to each mounting foot  18 . The ATMs  20  are affixed to and supported by the airframe structure (not shown). As noted above, there is also an ATM  20  located between each mounting foot  18  and the aircraft structure  14  (shown in FIG. 1).  
         [0030]    The ATMs  20  are part of an active transmission mount system  22  which is designed to isolate the main gearbox  12  from the airframe  14  so that the acoustic and vibratory loads are minimized or completely attenuated before passage into the fuselage. It is important to note that in order to isolate the gearbox  12  from the airframe, all the applied vibratory loads are isolated so that the gearbox  12 , in effect, floats in a dynamic sense with respect to the airframe, but maintains its steady, static position relative to the airframe. The system  22  positions the gearbox  12  by monitoring the sensed signals from a plurality of sensors  24  that are positioned on or adjacent to the mounting feet  18  and the ATMs  20 . The sensors  24  preferably sense parameters which can be utilized for determining the vibrational state of the mounting feet  18  (or in the structure adjacent to the mounting feet  18 ) and for determining the operational state, e.g., actuation position or oscillation mode, of the ATMs  20 . In one embodiment of the invention, position sensors  24  monitor the steady position of the feet relative to the airframe. A controller  26  operates to nullify the position off-set. Concomitantly, vibration sensors  25  sense the airframe vibrations and output a corresponding signal. This signal is used by the controller  26  to provide the vibratory flow to the ATM to nullify the vibration. The vibration sensors can be replaced with pressure sensors placed in the fluid lines  40 . In this case, the controller  26  nullifies the vibratory pressure, thus reducing the vibratory load passing through the ATM&#39;s into the airframe. This results in lower fuselage vibration. The figure illustrates the use of position sensors  24  and accelerometers  25  mounted at strategic locations within the airframe  14  to sense accelerations for feedback to the system for processing.  
         [0031]    As discussed above, the sensed signals from the sensors  24  and  25  are provided to an electronic controller  26 , such as a signal processor or computer. The controller  26  determines the position and loading being transmitted from each mounting foot  18 . The controller  26  then determines a desired operational state for each ATM  20  as a function of one or more of the sensed signals  24  and  25 .  
         [0032]    The controller  26  sends a signal to an actuation system  28  commanding the actuation system to actuate the ATMs  20  according to the desired operational state. In one embodiment of the invention, the actuation system  28  is a hydraulic actuation system and the ATMs  20  are hydraulic actuators. A hydraulic actuation system is preferred since it permits high frequency and high pressure control of the actuators to accommodate the frequencies and high loads typically experienced in a helicopter aircraft system.  
         [0033]    The hydraulic actuation system  28  supplies hydraulic fluid under pressure to each hydraulic actuator  20  to control the actuator so that it moves in the desired direction and at the desired frequency to counteract the sensed vibrations. In the illustrated embodiment, the hydraulic actuation system  28  preferably includes one or more electro-hydraulic valves  30  which are each electrically connected to the controller  26  via a control line  32 . The control line  32  is used to supply current to the valve  30 . The valve  30  controls flow of high pressure hydraulic fluid from a fluid source (not shown) to the ATM actuators  20 . The supply flow into the valve  30  is indicated by the numeral  34  and the flow to the ATM actuators  20  is indicated by the numeral  36 .  
         [0034]    The actuation system  28  also preferably includes a tuned acoustic stub assembly  38 . (Note: for the sake of simplicity, the fluid lines  40  are shown going to only two of the actuators.) The tuned stubs introduce softness into the hydraulic system at pre-selected frequencies to allow the system to attenuate high frequency (e.g., 778 Hz, 1556 Hz) and low amplitude (e.g., {fraction (1/1000)} inch) acoustic vibrations that are otherwise transmitted by the gearbox feet  18  into the ATM, thus, causing high frequency pressure in the fluid lines. It is also contemplated that an accumulator can be substituted for the tuned stubs to assist in pressurizing the various ATM actuators and adding softness into the system.  
         [0035]    The ATM actuator  20  according to the present invention isolates the gearbox from the airframe and, thus, controls the transmittance of vibratory loads. The ATM actuator  20  is designed to accommodate the large quasi-steady flight loads that are transmitted through the mounting feet  18  to the airframe. The magnitude of these loads can be quite high. For example, in a Sikorsky Aircraft Corporation S-76 aircraft, the quasi-steady flight loads have a magnitude of about 8000 lbs. (which is the approximate weight of the aircraft) on each actuator  20 . This applied quasi-steady flight load can change in magnitude and direction very quickly depending on pilot inputs into the aircraft. As such, the ATM actuator  20  must be designed to accommodate such loads while limiting gearbox motions to only about ±0.050 inches in order to avoid excessive misalignment of the engine-transmission shaft. The actuator according to the present invention is also designed to prevent the transmission to the airframe of small vibratory loads (e.g., 500 lbs. between about 16 Hz and about 50 Hz) and even smaller acoustic loads (30 lbs. between 600 Hz and 3 kHz) which are the vibratory and acoustic loads which cause the vibration and acoustic noise that are the most bothersome to the passengers and crew within the aircraft.  
         [0036]    The actuator  20  is designed to passively isolate the vibratory and acoustic loads which are applied perpendicular to the longitudinal axis of the actuator. This is achieved by configuring the actuators so that the off-axis (perpendicularly applied) loads are forced to pass through the soft axis of the actuator&#39;s built-in elastomeric element (discussed in greater detail hereinbelow). The loads that are applied along the actuator longitudinal axis are transmitted into the hydraulic column. This causes pressure fluctuations which would otherwise be transmitted into the airframe causing noise and vibration. However, the pressure is altered by the electro-hydraulic valve  30  (shown in FIG. 2) to minimize either vibratory pressure, thereby reducing the transfer of low frequency vibrations into the airframe. High frequency pressures that would cause noise in the aircraft are attenuated by the tuned acoustic stubs  38 .  
         [0037]    Referring now to FIGS. 3 and 4, one preferred hydraulic ATM actuator  20  is shown which meets these design requirements. The actuator  20  is a lap-fit type actuator and includes a housing  44  with a mounting flange  42  for attaching the actuator  20  to the airframe and/or a support structure. The actuator  20  also includes a mounting member  46 , such as a threaded stud, that is configured to attach to the mounting foot (not shown). Other methods of attachment may be used with the present actuator design.  
         [0038]    The mounting member  46  is attached to a piston  48  which is adapted to slide within a sleeve  50 . As discussed above, the actuator  20  is preferably a lap-fit type actuator. Lap-fit actuators are well known in the art. These actuators do not include any sliding seals. Instead, the lap-fit actuator is designed with extremely close tolerances (i.e., millionths of an inch) between the piston and the outer sleeve. As such, there is a close sliding engagement between the outer surface of the piston  48  and the cylindrical inner surface of the sleeve  50 . Seals have two key drawbacks. First, they eventually wear out under the action of continuous vibration. Second, the seals provide an undesirable path for vibratory and acoustic loads in the housing  44  and, thus, the airframe  14 . Loads transmitted through this path cannot be easily controlled by the valve  30  or attenuated by the tuned stub  38 . Those skilled in the art of actuators are well aware of the design specifics of a lap-fit type piston arrangement and, therefore, no further details are needed. Moog Inc. is one of several manufacturers currently making lap-fit type actuators.  
         [0039]    Since manufacturing a piston  48  to meet the tolerances needed for a lap-fit is difficult, the outer surface of the piston  48  in the present embodiment is formed with two raised annular surfaces  48   A  which provide the close sliding engagement between the piston  48  and the sleeve  50 . The sleeve  50  is located within a recess of the housing  44 . A static seal  52 , such as an O-ring seal is located between the sleeve  50  and the housing  44  to prevent hydraulic fluid from passing between the two.  
         [0040]    As discussed above, the present invention is designed to isolate the vibratory and acoustic loads from being transmitted from the gearbox feet  18  to the airframe. This is accomplished by canceling or attenuating the vibratory or acoustic loads that would otherwise pass into the airframe  14  by either actively oscillating the flow into the actuator  20  (or more exactly the hydraulic space  86 ) using the valve  30 , or passively attenuating acoustic pressure through the added softness of the tuned acoustic stubs  38 . However, quasi-steady loads which change depending upon the pilot inputs and which approximately represent the aircraft weight amplified by the maneuver the aircraft is undergoing, must be transmitted to the airframe  14 .  
         [0041]    Hence, the actuator  20  must be designed to accommodate these quasi-steady loads. Referring to FIG. 4, the load applied by the gearbox foot  18  to the stud  46  has a vertical quasi-steady state load V which is subject to a horizontal motion H. The vertical load V on a typical S-76 aircraft manufactured by the Sikorsky Aircraft Corporation is approximately 8000 lbs and the horizontal motion H is approximately ±0.050 inches and can be either steady or vibratory in nature. The maximum horizontal load is approximately 1500 lbs. There is also some vertical motion that occurs due to loading on the transmission foot. This vertical motion is approximately 0.075 inches from the steady state loading and 0.050 inches from the vibratory loading. These motions result in a shear loading and a moment loading on the actuator  20 . As discussed above, the present invention includes a lap-fit type piston which cannot withstand side loads higher than about 50 pounds. As such, a lap-fit type piston could not be used in a conventional actuator to react these applied loads. The present invention, however, utilizes a novel bearing assembly  54  to react the applied loads. The bearing assembly  54  is described in more detail below.  
         [0042]    The bearing assembly  54  is located on top of the sleeve  50  and includes an outer bearing member  56  which contacts an upper edge of the sleeve  50 . The outer bearing member  56  is attached to the housing  44  through any conventional method. In the illustrated embodiment, a clamping ring  58  is used to attach the outer bearing member  56  to the housing  44 . More particularly, the clamping ring  58  contacts a flange  60  on the outer bearing member  56 . The flange  60 , in turn, presses against a seal  62  that is adjacent to the housing  44 . The clamping ring  58  is attached to the housing  44  with conventional fasteners (not shown) and, thus, locks the bearing assembly  54  and sleeve  50  to the housing  44 .  
         [0043]    In one embodiment of the invention, the housing  44  and clamping ring  58  are made from stainless steel material, although aluminum is preferable. The mounting member  42  is preferably made from aluminum material. The piston  48  and sleeve  50  are preferably made from stainless steel.  
         [0044]    A diaphragm  64  is attached to a lower face of the outer bearing member  56  and separates the bearing assembly  54  from the sleeve  50  and the piston  48 . As such, slightly pressurized hydraulic fluid which leaks past the piston  48  is substantially prevented from passing from the sleeve/piston side of the diaphragm  64  to the bearing assembly side. The diaphragm  64  is preferably made from nitral rubber. This is done because hydraulic fluids can damage the elastomeric material of the outer bearing member  56 . The space between the diaphragm  64  and the other components of the outer bearing member  56  is preferably filled with polybutene. This material does not react unfavorably with the elastomer of the outer bearing member  56  when in direct contact. The material prevents excessive deformation of the diaphragm  64  when subjected to the low pressure caused by leakage of hydraulic fluid past the lap-fit piston  48 . The diaphragm  64  is preferably attached on its inner diameter to the bearing assembly  54  at point  88  and/or to the piston rod  78 . The diaphragm  64  is preferably attached at its outer diameter to the outer bearing member  60  at point  89 .  
         [0045]    The bearing assembly  54  is designed to react the loads that are applied to the actuator  20  from the gearbox foot  18 . The bearing assembly  54  includes three resilient bearings to react or isolate the applied loads to effectively eliminate any side loading on the piston  48  and sleeve  50 . Referring to FIG. 4, a first or thrust bearing  66  is mounted below the stud  46  and is preferably an elastomeric bearing. Elastomeric bearings are well known in the art and generally comprise alternating layers of elastomer and nonresilient shims. The number of elastomer layers and shims is not limited to the number shown in the figures but, instead, would be determined by the applied loads. The thrust bearing  66  is designed to be stiff (i.e., rigid) in the axial (vertical) direction and soft (i.e., flexible) in the lateral (horizontal) direction. For a 12,000 lb gross weight helicopter, the axial stiffness of the thrust bearing  66  is preferably greater than about 900,000 lbf/in, and the lateral stiffness is preferably less than about 2000 lbf/in. The thrust bearing  66  is preferably substantially planar in shape. The thrust bearing  66  is preferably located between and attached to a mounting member formed on the stub  46  and a thrust mounting plate  68 . The attachment of the bearing is through any conventional means, such as adhesive.  
         [0046]    The thrust plate  68  is attached to or formed on an upper end of an inner bearing member or support  70 . The inner bearing member  70  is preferably substantially cylindrical in shape with a longitudinal axis that lies substantially in line with a longitudinal axis of the piston  48 . At least a portion of the inner bearing member  70  is located within the outer bearing member  56 . A radial journal bearing  72  is located between the inner and outer bearing members  56 ,  70 . The radial journal bearing  72  is preferably an elastomeric bearing that is cylindrical in shape. The radial journal bearing  72  is preferably stiff in the radial direction and soft axially. For a 12,000 lb. gross weight helicopter, the radial journal bearing preferably has a radial stiffness greater than about 300,000 lbf/in and an axial stiffness less than about 4000 lbf/in.  
         [0047]    A spherical bearing  74  is located within and engages with an inner spherical surface on the inner bearing member  70 . The spherical bearing  74  is attached to a spherical bearing support  76 . The spherical bearing  74  is preferably an elastomeric bearing with a center line that lies substantially in line with the longitudinal axis of the piston  48 . The spherical bearing  74  is preferably stiff axially and soft tangentially. For a 12,000 lb. gross weight helicopter, the spherical bearing preferably has an axial stiffness greater than about 500,000 lbf/in, and a tangential stiffness less than about 2320 in-lbf/deg. The center of rotation of the spherical bearing  74  is preferably close to the center of rotation of the radial journal bearing  72 . In this way, any small angular motions of the inner bearing member  70  does not produce large moments or radial loads on the spherical bearing support  76 .  
         [0048]    The cumulative bearing stiffness (without hydraulic fluid) assures isolation of vibratory and acoustic motions in the axial direction from being transmitted into the mounting flange  42  and, thus, into the airframe  14 . For a 12,000 lb. helicopter, the cumulative bearing stiffness in the horizontal and vertical directions is preferably less than about 20,000 lbf/in. The cumulative axial stiffness of the thrust bearing  66  and the spherical bearing  74  must be as high as possible, preferably, greater than 100,000 lbf/in, to preclude excessive compression of these members under vibratory load. This compression causes unwanted flow demands upon the hydraulic valve  30  and the supply  34 .  
         [0049]    The spherical bearing support  76  is engaged with the piston  48  through a piston rod or stinger  78 . The piston rod  78  is preferably made from a metallic material, such as steel and is attached to the spherical bearing support  76  and piston  48  through any conventional means known to those skilled in the art, such as a bolted connection  80 . In one embodiment of the invention, the piston rod  78  is designed to have some lateral flexibility (i.e., reduced lateral stiffness) to permit bending and, thus, further eliminate any residual side loads from being transmitted to the piston  48  from the bearing assembly  54 .  
         [0050]    The preferred bearing assembly  54  provides a novel mechanism for transferring and reacting applied gearbox foot  18  loads with a lap-fit type actuator. As stated above, the applied loads result in moment and shear loads being applied to the stud  46 . The design of the bearing assembly  54  is such that axial loads are transmitted through the thrust bearing  66  and spherical bearing  74  directly to the piston  48 . The lateral loads caused by the lateral motions are reduced by radial shearing of the thrust bearing  66 . The residual loads are reacted by the radial journal bearing  72  which prevents the inner bearing member  70  from radial motion and cocking with respect to the outer bearing member  56 . Any residual cocking of the inner bearing member  70  is accommodated by the spherical bearing  76  which is very soft circumferentially. As a consequence, only small angular moments are transferred through the spherical bearing  74  to the spherical bearing support  76  and, thus, to the piston rod  78 . The lower end of the piston rod  78  reacts these small moments as a small radial force. A slip surface (not shown) can be formed at the lower end of the piston rod  78  to eliminate this small side load if desired or, instead, the piston rod  78  can be designed to bend.  
         [0051]    The present invention acts as an isolation system for preventing vibration and noise from transferring from the gearbox feet  18  to the airframe. The active transmission mount system  22  applies a quasi-steady pressure to the actuator  20  to react the applied vertical steady state loads. The system then controls the flow within the actuator  20  to relieve pressure when a vibratory load pushes on the actuator  20  and increase the flow when the vibratory load pulls on the actuator. Hence, the actuator  20  is operated by removing and supplying a sufficient amount of hydraulic fluid against the head of the piston  48  to translate the piston in substantially the same direction and at substantially the same frequency as the applied vibratory loading. As a result, vibratory pressure is minimized and, thus, the actuator  20  does not transmit any vibration or noise related loads to the airframe structure. The applied flows are controlled by the controller  26  and the hydraulic actuation system  28 .  
         [0052]    The pressurized hydraulic fluid is channeled through an inlet  84  to a chamber  86  and against the head of the piston  48 . Since a lap-fit piston does not include seals, pressurization is obtained by a close-fit sliding interface between the piston  48  and the sleeve  50 . However, even with this close fit, hydraulic fluid will escape between the piston  48  and the sleeve  50  as the piston  48  moves relative to the sleeve  50 . To account for the leakage, the present invention includes a outlet passage  87  which channels the low pressure leaking hydraulic fluid back to the hydraulic actuation system  28 .  
         [0053]    The present invention provides a novel actuator and system for controlling vibrational and acoustic noise transmission from a helicopter gearbox footing to the airframe  14 . While the invention is illustrated and described as being used for controlling vibrational transmission from a gearbox foot to the airframe  14 , it is not limited to that embodiment. On the contrary, the present invention can be used to address vibrational transmissions from a variety of other components in various types of machines and aircraft.  
         [0054]    An ATM actuator  20  and ATM system  10  were tested on the gearbox footing of a Sikorsky Aircraft Corporation S-76 aircraft. Reductions of about 13 dB in cabin noise and about 20 db in vibrations have be achieved using the present invention.  
         [0055]    Although the invention has been described and illustrated with respect to the exemplary embodiments thereof, it should be understood by those skilled in the art that the foregoing and various other changes, omissions and additions may be made therein and thereto, without parting from the spirit and scope of the present invention.