Patent Abstract:
Disclosed herein are control systems relating generally to the field of aerodynamics and more particularly to the control of vibration of rotor blades such as helicopter blades. Such systems involve devices for vibration control of each rotor blade, which incorporate control systems of the flow control type (e.g. actively controlled flap) and structural control type (e.g. active pitch link). Also disclosed are related methods of controlling vibration in a rotor blade, wherein the rotor blade is coupled to a rotor hub and has at least a torsional stiffness and a pitch angle associated therewith.

Full Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     This application claims the priority right of prior U.S. patent application Ser. No. 61/076,171 filed on Jun. 27, 2008 by applicants herein. 
    
    
     TECHNICAL FIELD 
     The invention relates generally to the field of aerodynamics and more particularly to a system and method for the control of vibration of helicopter rotor blades. 
     BACKGROUND TO THE INVENTION 
     As those skilled in the art are aware, both flow control and structural control devices can be employed on each rotating rotor blade of a helicopter to minimize vibration in flight. The most efficient method of reducing vibration on helicopter rotor blades is through Individual Blade Control (IBC) in which each rotor blade is individually controlled using a flow control or structural control device. 
     Structural control includes any devices capable of controlling the mass, stiffness or damping of the helicopter blade. The only practical structural control device developed to date is the Active Pitch Link, which is able to control the torsional stiffness characteristics of a blade. 
     Flow control can be defined as any control technique capable of controlling the aerodynamic loads acting on the blade. Such techniques include Actively Controlled Flap (ACF), Active Twist Rotor (ATR), Actively Controlled Tip (ACT), along with various types of Boundary Layer Suction/Blowing devices. For helicopters, the two most popular techniques have been the Actively Controlled Flap (ACF) and Active Twist Rotor (ATR). 
     There are a number of major research teams worldwide investigating the feasibility of various active control technologies on helicopter rotor blades. Of the research presently being performed, all research teams consider only one control system per blade. The most popular vibration control systems are of the flow control type with the most popular control system in this category being ACF because of the significantly lower power requirement than ATR. Some prior art systems have applied ACF with two independently controlled flaps on a single blade i.e. two independent control systems of the same type. 
     However, the problem with applying only one type of control device, especially actively controlled flap (ACF) or active twist rotor (ATR), is that these devices are not very efficient on their own. This is due to the fact that both of these technologies try to actively control the twist (or effective pitch angle) of the rotor blades. This is clearly the goal of a rotor blade employing ATR, but even with ACF it has been shown that a flap is much more efficient when used as a servo-tab than when used as a high-lift device. The goal of a servo-tab is to twist the rotor blade as a result of the flap deflection whereas the goal of the high-lift device is to increase the local rotor blade section lift of a rigid blade. 
     In order to impose the highest possible twist effect, either as a result of employing ACF or ATR technology, the rotor blade torsional stiffness should be as low as possible. However, the torsional stiffness of a helicopter rotor blade is set to a certain level to avoid excessive deformations due to the aerodynamic loads during operation. This level cannot be lowered by simply making softer blades; otherwise the blades would become too flexible and aeroelastic problems and loss of aerodynamic efficiency would occur. 
     Therefore, there is a need in the art for some kind of control system allowing the rotor blade torsional stiffness to be lowered whenever the flow control device is actuated. 
     SUMMARY OF THE INVENTION 
     Certain exemplary embodiments may provide a feedback control system for controlling vibration in a rotor blade, wherein the rotor blade is coupled to a rotor hub and has at least a torsional stiffness and a pitch angle associated therewith, the feedback control system comprising: a flow control device for adjusting the pitch angle of the rotor blade; a structural control device for adjusting the torsional stiffness of the rotor blade; a plurality of sensors attached to the rotor blade; and a control computer communicating with the flow control device, the structural control device and the plurality of sensors, wherein vibration data from the sensors is received by the control computer and control signals are generated by the control computer to reduce the torsional stiffness of the rotor blade with the structural control device and simultaneously increase the pitch angle of the rotor blade with the flow control device. 
     Certain other exemplary embodiments may provide a method of controlling vibration in a rotor blade, wherein the rotor blade is coupled to a rotor hub and has at least a torsional stiffness and a pitch angle associated therewith, the method comprising the steps of: receiving vibration data from a plurality of sensors into a control computer, wherein the control computer communicates with a flow control device, a structural control device and the plurality of sensors, wherein each of the flow control device, the structural control device and the plurality of sensors are electromechanically coupled to the rotor blade; generating control signals in the control computer; adjusting the structural control device to reduce the torsional stiffness of the rotor blade based on the control signals inputted therein; and simultaneously adjusting the flow control device to increase the pitch angle of the rotor blade based on the control signals inputted therein. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention will now be described with reference to the drawings in which: 
         FIG. 1A  depicts an overview of the rotor blade incorporating the hybrid vibration control device of the present invention; 
         FIG. 1B  depicts the sensors integral to the hybrid device of  FIG. 1A ; 
         FIG. 2A  depicts the main parts of a prior art helicopter rotor hub; 
         FIG. 2B  depicts a collective change of the pitch angle of the rotor blades of  FIG. 2A ; 
         FIG. 2C  depicts a cyclical change of the pitch angle of the rotor blades of  FIG. 2A ; 
         FIG. 3A  depicts a prior art smart spring; 
         FIG. 3B  depicts the smart spring of  FIG. 3A  with the actuator off; 
         FIG. 3C  depicts the smart spring of  FIG. 3A  with the actuator on; 
         FIG. 4A  depicts schematically the smart spring used in an active pitch link (APL) in the hybrid device of the present invention; 
         FIG. 4B  depicts the primary (fail safe) load path of the smart spring of  FIG. 4A ; 
         FIG. 4C  depicts the secondary load path of the smart spring of  FIG. 4A ; 
         FIG. 4D  depicts a perspective view of the active pitch link (APL) in accordance with the present invention. 
         FIG. 4E  depicts an exploded view of the APL in accordance with the present invention. 
         FIG. 4F  depicts graphically the damping characteristics of the APL of  FIG. 4A ; 
         FIG. 4G  depicts the theoretical modes of operation of the APL of  FIG. 4A ; 
         FIG. 4H  depicts graphically the actual modes of operation of the APL of  FIG. 4A ; 
         FIG. 5A  depicts an active control flap operating in the high-lift mode; 
         FIG. 5B  depicts an active control flap operating in the servo-tab mode; 
         FIG. 5C  depicts an active control flap in accordance with the present invention; 
         FIG. 5D  depicts a side view of the active control flap of  FIG. 5C  with the piezoelectric actuators in the “off” and “on” positions; 
         FIG. 5E  depicts the skeleton and frame to which the active control flap of  FIG. 5C  is attached; 
         FIG. 5F  depicts a fully assembled rotor with the active control flap of  FIG. 5C ; 
         FIG. 6A  depicts the fan plot for a helicopter blade and the effect of pitch link stiffness on torsional mode frequency; 
         FIG. 6B  depicts an experimental demonstration of centrifugal tests to show the reduction of torsional stiffness via altering the resultant stiffness of the pitch link. 
         FIG. 7  is a block diagram depicting the control system of the hybrid device of the present invention; 
         FIG. 8  depicts a flow chart detailing the control steps performed by the control computer of  FIG. 7 ; 
     
    
    
     DESCRIPTION OF PREFERRED EMBODIMENTS 
     (A) Hybrid Device—Overview 
     The present invention employs, at least in selected embodiment, the simultaneous application of any type of structural control and flow control device on each individual blade. For the purposes of describing the invention, a specific example of employing an Active Pitch Link (APL) for structural control and Actively Controlled Flap (ACF) for flow control will be presented. However, it should be appreciated, that the invention is not meant to be limited to this embodiment. The general principle of combining these two devices in a “hybrid system” can be extended to any other combination of structural and flow control devices. 
     Referring to  FIG. 1A , the structural control is realized using an Active Pitch Link (APL)  100 , which replaces the conventional pitch link on the rotor hub ( 102 ). APL  100  is capable of changing the torsional stiffness of the rotor blade  104 . 
     The flow control is realized via an Actively Controlled Flap (ACF)  106 , located at the trailing edge of rotor blade  104 , closer towards the tip. 
     The frequency at which these two mechanisms operate is important. Both are able to actuate at the frequencies typical of Individual Blade Control (IBC), i.e. between (N−1)/rev and (N+1)/rev, where N represents the number of rotor blades  104 , i.e. for a 4-bladed rotor, both systems should have the capability to operate at the frequency of 3 to 5 actuations per revolution. 
     The two systems are connected to a controller  108  located on the top of rotor hub  102 , which dictates the combined operation of the two systems with the goal of minimizing vibrations. As depicted in  FIG. 1B , the entire system is equipped with preferably eight (8) sensors  110  measuring vibration. Sensors  110  include a strain gauge, two hall sensors and three accelerometers mounted on rotor blade  104  and one hall sensor and two accelerometers mounted on APL  100 . Sensors  110  are linked to a computer in controller  108 , thus forming a closed-loop feedback control system consisting of controller  108 , APL  100 , ACF  106  and sensors  110 . The feedback control system will be discussed in more detail in relation to  FIGS. 6 and 7 . 
     (B) Structural Control—Active Pitch Link (APL) 
     The main parts of a typical helicopter rotor hub  200  are depicted in  FIG. 2A , highlighting the location of swashplate  210  as well as conventional pitch link  220 . It is the lower non-rotating disk of swashplate  210 , which is controlled by the pilot (not shown). When swashplate  210  is moved up-down or tilted to any direction, upper rotating disk  230  follows swashplate  210  and this motion is transferred to the rotor blade  240  via pitch link  220 . The purpose of the swashplate-pitch link system is to change the pitch angle θ of rotor blade  240  and thereby the magnitude and tilt of the resultant thrust force generated by rotor hub  200 . The pitch angle θ of rotor blade  240  can be changed either collectively (via an up-down motion of swashplate  210  (as depicted in  FIG. 2B ) or cyclically via tilting of swashplate  210  (as depicted in  FIG. 2C  or in any combination of both. Each rotor blade  240  is connected to swashplate  210  via an associated pitch link  220  and pitch horn  250 . Pitch horn  250  is essentially the moment arm of pitch link  220 , allowing the rotation (“pitching”) of rotor blade  240  along its longitudinal (spanwise) axis. Changing the stiffness of rotor blade  240  at the root requires some form of active control system located at the root of rotor blade  240 , either directly at the root section of rotor blade  240  or indirectly on rotor hub  200 . 
     Active Pitch Link—Operating Principle 
     The Active Pitch Link (APL) of the present invention, at least in some embodiments, is a piezoelectric actuator-based device for controlling the blade stiffness at the root. The APL replaces conventional pitch link  220  on rotor hub  200 . Thus, its primary purpose is to control the pitch angle of rotor blade  240  in a semi-active way. The term semi-active control is used since the APL utilizes the concept of a Smart Spring as described in U.S. Pat. No. 5,973,440 entitled “Structural Component Having Means for Actively Varying its Stiffness to Control Vibrations”, issued Oct. 26, 1999 to Nitzsche et al. which is incorporated by reference herein. The described Smart Spring allows a user to control the displacement of a device in one direction only—the direction in which the load acts on the device. A fully-active control system would allow displacements in both directions, i.e. also in the direction opposite to the force acting on the device. 
     The operational principle of a generic Smart Spring is shown in  FIG. 3A . Two springs, k 1  and k 2  have their ends attached to opposing walls  300  and a pair of sleeves  310 ,  312  that can slide one with respect to the other. An external (input) force F is applied to sleeve  312 . A stack of piezoelectric actuators  320  is inserted into sleeve  310 . 
     Referring to  FIG. 3B , when the actuator is “OFF”, the sleeves  310 ,  312  can move freely and the resulting horizontal displacement (output) is δ max =F/k 2 . Spring k 2  is designed to be the “primary” load path of the APL. Referring to  FIG. 3C , when the actuator is turned “ON”, sleeve  310 , under the action of the stack of piezoelectric actuators  320 , yields and applies on sleeve  312   a  resultant normal force, N. 
     A friction force, μN is induced by the contact between the surfaces of sleeves  310 ,  312 . If this friction force is sufficiently large and sleeves  310 ,  312  are forced into motion together, springs k 1  and k 2  act in series and a smaller horizontal output displacement δ min =F/(k 1 +k 2 ) is obtained because the stiffness experienced by the input force rises from the system&#39;s original k 2  to k 1 +k 2 . Spring k 1  is driven by the resultant friction force μN applied by the sleeve  310  on sleeve  312 , which is controlled by the external electrical stimulus (control input) to the stack of piezoelectric actuators  320 . Spring k 1  is called the. “secondary” path of the APL. 
     Thus, the horizontal output displacement of the system under the input force F varies between the referred two extremes, F/(k 1 +k 2 )≦δ≦F/k 2  and the total load is distributed between the primary and the secondary load paths. 
     The APL system also changes its apparent mass because the stack of piezoelectric actuators  320  and sleeve  310  have inertial properties. However, this effect can be disregarded if the overall system is “stiffness dominated” (i.e., the harmonic disturbance force has a frequency much lower than the internal resonance frequencies of the APL). The dry friction between sleeves  310 ,  312  also creates coulomb damping, which cannot be neglected. The latter adds an important stabilizing effect to the system. Since the APL actively changes both its apparent mass and stiffness and also its internal damping, it is called an “impedance control” device. 
     As discussed above, within the context of helicopter applications, the active pitch link (APL) replaces conventional pitch link  220 . Thus, rotor blade  240  and the APL become an integral system, which can control the twist impedance of rotor blade  240  in real time, by targeting the 1st torsional mode of rotor blade  240 . However, because of the inherent coupling between blade modes (i.e. when a blade is twisted, it will generate more lift, i.e. it will bend/flap up and as a result of this motion it will generate lead-lag motion too), when the torsional mode is controlled, all modes are controlled. 
     Active Pitch Link—Design 
     Referring to  FIG. 4A , although the APL  400  of the present invention uses the Smart Spring concept, its internal configuration is significantly altered to facilitate a feature very important for aerospace applications: fail safe design. Fail safe design means that when a power failure or failure of piezoelectric actuator  410  occurs, APL  400  returns to the original “conventional pitch link” mode. In order to fulfill this fail safe design requirement, springs k 1  and k 2  are incorporated in parallel rather than in series (as in  FIG. 3A ). Using this configuration, the overall system stiffness can be varied between k 1  (“soft” link) and k 2  (“solid” link), instead of the ranges of k 2  and k 1 +k 2 . 
     The main parts and operation of the APL  400  are arranged in the following configuration. In the default position i.e. when piezoelectric actuator  410  is OFF, a preload spring  420  pushes a friction pad  430  to a pair of solid links  440   a  and  440   b . The force generated by preload spring  420  is such so that the friction force between friction pad  430  and solid links  440   a ,  440   b  is larger than the overall vertical force acting on APL  400 . Thus, when piezoelectric actuator  410  is OFF, all of the load will be transferred from a top plate  450  to a bottom plate  460  via load path consisting of solid link  440   a , friction pad  430  and solid link  440   b.    
     When piezoelectric actuator  410  is ON, friction pad  430  is pushed away from the solid links  440   a ,  440   b  and, when the two surfaces disengage, the entire load is transferred from top plate  450  to bottom plate  460  via “soft” spring k 1 . 
     An intermediate mode of operation, called transitional mode, can also be generated. This occurs when piezoelectric actuator  410  is only partially activated (i.e. when the actuation power is somewhere between zero and the maximum voltage). In this case, sliding friction will occur between friction pad  430  and solid links  440   a ,  440   b , thus initiating the “energy extraction” operational mode, in which vibration is reduced by extracting energy from the system via sliding friction and heat. 
     The operational principle of APL  400  are illustrated in  FIGS. 4B and 4C  which depict the load paths when the actuator is switched ON and OFF. 
     A more detailed depiction of APL  400  is provided in  FIGS. 4D and 4E . Here, the two springs k 1 , k 2  shown previously in the schematic diagram of  FIG. 4A  are arranged in a concentric fashion, i.e. solid link k 2    461  slides into the soft link spring k 1    462 . Such arrangement enhances the compactness of the design, which is important because of the space limitations on a rotor hub. 
     The two cylindrical piezoelectric actuators  464  are held in a holder assembly  466 , including friction pad  468 , preload springs  470 , shoulder bolt  472 , load cell  474  and a pair of set screws  476 . 
     Piezoelectric actuators  464  are off-the-shelf units from Piezomechanik Gmbh, capable of generating 1800 N block force or 60μ of displacement. Friction pad  468  is made out of brass, an effective material from friction point of view. The preload spring  470  is realized via a set of wave disc springs, which offer modularity (their number can be varied) as well as compactness. The amount of preload force can be adjusted via the number of wave disc springs applied as well as via the 2 set screws  476 . The resultant force acting on the friction pad (i.e. the sum of the preload spring force and the actuation force) is monitored via load cell  474 . Washers  478 ,  480 , spacer  482 , screw  484  and nuts  486  and  488  all serve to hold the whole holder assembly together. 
     There are two discs  490  mounted on the top and bottom of APL  400  which hold accelerometers  492 , measuring both the vibratory loads as well as the relative displacement of the upper and lower swivel joints  494 . The top swivel joint  494  is left threaded and connects to the pitch horn of the rotor blade, whereas the lower swivel joint  494  is right-threaded and connects to the swashplate. Nuts  495  counter swivel joints  494  and thus serve to adjust the length of APL  400 . 
     Custom screw  496  serves to connect shoulder bolt  472  to solid link  461 . This is required to ensure that the friction force generated by piezoelectric actuators  464  is independent of the centrifugal loads, which should act from the load cell  474  towards the piezoelectric actuators  464 . 
     In addition to accelerometers  492 , the performance of APL  400  can also be monitored via a built-in Hall sensor  498 . Hall sensor  498 , mounted on soft spring link  462 , is paired up with a permanent magnet  499 , mounted on friction pad  468 . As these two move relative to each other, the electrical signal in the Hall sensor  498  changes and this can be related to the displacement between the two parts. The exact location of the permanent magnet  499  is adjustable since it is threaded at the bottom. 
     The APL depicted in  FIGS. 4D and 4E  operates as follows. As a default, the piezoelectric actuators  464  are OFF and preload spring  470  pushes friction pad  468  to the side of soft link spring  462 . The preload force has to be set in a way so that the default friction force is large enough to overcome the vertical force acting on APL  400 . Thus, solid link  461  and the top of soft spring link  462  become locked via the friction pad  468  (i.e. they cannot move relative to each other) and the load acting on APL  400  will be transferred from top to bottom via the following path: top swivel joint  494 —top of soft spring link  462 —friction pad  468  (link via friction)—custom shoulder bolt  472 —solid link  461  (connection via screw  496 )—bottom swivel joint  494 . 
     When the piezoelectric actuators are ON, friction pad  468  slides on custom shoulder bolt  472  and disengages the friction pad  468  from the soft spring link  462 . Thus, soft spring link  462  and solid link  461  can move relative to each other since there is no link (via friction) between them. As a result, all vertical load acting on APL  400  will be transferred via the soft link spring  462  through the following load path: top swivel joint  494 —soft spring link  462 —bottom of solid link  461  (connection via thread)—bottom swivel joint  494 . 
     When the actuator is OFF, APL  400  operates in the solid link mode, thus providing a Fail Safe design. 
     The advantages of APL  400  of the present invention are numerous and include:
         (a) piezoelectric actuator  410  is used to generate friction force instead of acting against the principal force, thus requiring significantly lower power consumption (3-5% of Active Twist Rotor);   (b) APL  400  incorporates a fail safe design such that when a power failure or failure of piezoelectric actuator  410  occurs, loads are transferred via spring k 2  representing the “solid link”;   (c) the friction force generated through piezoelectric actuator  410  is independent of centrifugal loads. The system does not therefore lock purely from centrifugal loads;   (d) the system provides adjustable resultant system stiffness i.e. by careful adjustment of the actuator voltage, sliding friction can be generated between springs k 1  and k 2 . The sliding friction allows adjustment of the resultant system stiffness anywhere between k 1  and k 2  as depicted in  FIGS. 4F to 4H ; and   (e) the system allows for self-compensation due to wear i.e. if the damping characteristics of APL  400  change in time due to the wear of contacting parts or temperature increase, the control algorithm (discussed in relation to  FIG. 8 ) is able to self-compensate for these changes.
 
(C) Flow Control—Actively Controlled Flap (ACF)
       

     As will be understood by a skilled workman, an Actively Controlled Flap (ACF) can work in two modes: either as a) a high-lift device or b) as an aeroelastic servo-tab. As depicted in  FIG. 5A , the high lift device mode occurs when the blade behaves as a rigid structure, i.e. when the torsional stiffness of the rotor blade is very high. In this case, the local lift of the blade section is increased when the flap is deflected down. 
     As depicted in  FIG. 5B ), the aeroelastic servo-tab mode occurs when the blade behaves as an elastic structure, i.e. when the torsional stiffness of the blade is too low. In this case, the “soft” blade section rotates as a reaction to the flap deflection, i.e. the local lift of the blade section will increase when the flap is deflected upwards, in the opposite direction than before. However, this second mode can ultimately yield much higher overall blade lift than the first mode, because the angle of attack of the entire blade is increased in the servo-tab mode. In other words, if the blade is made “soft” enough in torsion, it can be essentially twisted up/down by activating the flap up/down, respectively. 
     It has been shown in the prior art that usually the servo-elastic tab mode is more effective for controlling vibration. Therefore, the operation of the present invention incorporates an Actively Controlled Flap (AOF) tailored to produce upward deflections only. 
     The design of the ACF of the present invention is depicted in  FIG. 5C . The ACF mechanism produces 4 degrees of deflection up (only) at a frequency of at least (N+1)/rev, where N is the number of rotor blades coupled to the rotor hub. 
     The proposed ACF  500  shown in  FIG. 5C  is driven by two piezoelectric actuators  505  which can operate at a frequency of up to 200 Hz. Hence, the system is capable of producing flap deflections corresponding to 8/rev for the worst case scenario of a scaled rotor with 1,555 RPM=25 Hz, i.e. well above the required (N+1)/rev (i.e. 5/rev for a 4-bladed rotor). The system is also capable of producing 4 degrees of deflection in the upward direction only. 
     As depicted in  FIGS. 5C and 5D , the basic principle of ACF  500  is that a sliding rod  510  connected to the actuators  505  slides back and forth. The rod end is connected to a wedge  515  which then slides on a moment arm  520  linked to the flap  525  via a hinge point  530 . As piezoelectric actuators  505  are activated, they increase their length and as a result sliding rod  510  moves forward (ΔX). At the same time, moment arm  520  moves down, thus rotating flap  525  up. Wedge  515  and moment arm  520  each contains a magnet  535 ,  540  of opposite poles which create a sliding link between the two parts. Magnets  535 ,  540  are sized in a way so that the two parts of moment arm  520  and sliding rod  510  do not lock. Note, however, that because helicopter blades typically operate at positive angles of attack, the aerodynamic force acting on the flap will always help to produce the upward deflection, whereas wedge  515  moving towards the trailing edge will push flap  525  down. 
     The flap system shown in  FIG. 5C  is attached to rotor blade  555  shown in  FIGS. 5E and 5F  via the attachment points  545  through a skeleton  550  (See  FIG. 5E ). Skeleton  550  is a removable part of rotor blade  555  which, during assembly, is slid into rotor blade  555  from the tip end. Skeleton  550  is a lightweight structure machined out of Titanium and optimized to bear stresses arising from the centrifugal loads of ACF  500 . ACF  500  slides into frame  560 , which is glued from inside to the skin of rotor blade  555 . Frame  560  features a nylon guiding rail for skeleton  550 . The two parts are connected to each other via a pin  565 , which is again sized to bear the resultant centrifugal loads from skeleton  550  and ACF  500 . The whole blade assembly is shown in  FIG. 5F . 
     (D) Operation of Hybrid Device 
     It has been shown in the prior art that vibration on helicopters can be reduced relatively successfully by imposing blade pitch angle changes of about 1 degree at a frequency ranging between (N−1)/rev to (N+1)/rev. It is for this reason that an Actively Controlled Flap (ACF) is preferred to be operated as an aeroelastic servo-tab instead of a high-lift device. 
     Achieving 1 degree pitch angle change, however, is at the limit of most flow control devices, such as the Active Twist Rotor (ATR) or Actively Controlled Flap (ACF). Larger pitch angle changes would lead to more significant reductions of vibration. The present invention allows the pitch angle change imposed by a flow control device to be improved by combining the flow control device with a structural control device. The structural control device serves to reduce the torsional stiffness of the blade whenever the flow control device is activated. More specifically, the blades are made instantaneously “softer” in torsion (twist) and thus the flow control device imposes larger pitch angles when activated. 
     This is the basic principle of the present invention, and a specific example would combine the Active Pitch Link (APL) (capable of controlling blade torsional stiffness) with the Actively Controlled Flap (ACF) (in the aeroelastic servo-tab mode) to create a “hybrid” control system. 
     The first condition of the hybrid control system is the careful selection of the stiffness of secondary “soft mode” spring k 1  of APL  400 . This value is selected in such a way that the natural frequency of rotor blade  240  in torsion, which is linked directly to torsional stiffness, (typically in the range of 6/rev) is brought down to the actuation frequency of the flow control device, APL  400  in the specific case of this invention (3/rev to 5/rev). The selection procedure of the soft mode spring stiffness is depicted in  FIG. 6A . On the left hand side, the fan plot of a typical helicopter blade is shown, illustrating that the natural frequency in torsion occurs at about 6/rev frequency. On the right hand side, a graph showing the result of a sensitivity study is shown. It depicts the variation in natural frequency in torsion with the resultant pitch link stiffness of rotor blade  240 . If, for example, the torsional mode of rotor blade  240  is to be brought down to 3/rev frequency, the stiffness of soft mode spring k 1  should be 180 kN/m according to the graph. Using this method, when the ACF is activated at 3/rev frequency and at the same time the APL is also activated bringing down the torsional frequency of rotor blade  240  to 3/rev, rotor blade  240  will resonate in torsion and thus larger twist angles can be achieved. 
       FIG. 6B  shows an experimental demonstration of the above claim from centrifugal tests. Note that for these tests a different blade was used than that described in the above computational studies. This meant that the “Soft Link” mode was expected to be achieved at a different spring stiffness than in the computational results shown above. Various spring stiffnesses were tested, ranging from a practically infinite value (k 5 ˜2,000 kN/m), representing the “Solid Link” mode of the APL, to a very low one (k 1 =10.9 kN/m) representing the “Soft Link” mode. Intermediate spring values were also considered to represent the transitional mode, i.e., k 4 =160.0 kN/m, k 3 =82.7 kN/m. From the fan plots, it is evident that the first torsional mode is indeed affected by the variation of the resultant pitch link stiffness at all rotational speeds. As expected, the torsional stiffness decreases as the APL becomes “softer”. The magnitude of the change is viewed relatively small, which would call for even lower APL stiffness in future iterations. However, the concept of reducing torsional stiffness via altering the resultant stiffness of the pitch link is successfully demonstrated in these experiments. 
     A block diagram of the “hybrid” control system is depicted in  FIG. 7 . As highlighted in the figure, there is one central control computer  710  in the system, located preferably on the top of rotor hub  102  (See  FIG. 1A ). Control Computer  710  serves all N blades. In order to realize the Individual Blade Control (IBC) integral to the present invention, each rotor blade  104  (See  FIG. 1A ) has to be equipped with its own individual control system, i.e. each rotor blade  104  includes a structural control device (e.g. APL  400 )  720  and a flow control device (e.g. ACF  500 )  730  i.e. structural control device  720  and flow control device  730  will occur N times on helicopter rotor hub  102 . As shown in the figure, the control reference parameter (IN)  740  is the desired level of vibration. The actual level of vibration is measured via the eight (8) sensors  750  located on each rotor blade  104  (See element  110  in  FIG. 1B ). Sensors  750 , along with three (3) accelerometers  760  located on the rotor shaft (not shown in the  FIGS. 1A and 1B ) provide a feedback signal to control computer  710 , which then determines the optimum strategy for minimizing vibration and provides a control signal to structural control device  720  and flow control device  730 . 
       FIG. 8  depicts a flow chart detailing the control steps performed by control computer  710  of  FIG. 7 . First, vibration data is received from sensors  110  at step  805 . This data, along with a certain portion of the time history of previous data, are analyzed via Fourier transformation at step  810  to determine the dominant vibration frequency (f VIB ) and vibration amplitude (P VIB ). 
     Following this, the type of control strategy (i.e. “ACF only”, “APL only” or “hybrid” control) is determined at steps  820 ,  830  or  840  based on either the manual input of the pilot/operator or a database of experimental tests, in which the various control strategies have been linked to certain vibration levels. 
     Starting from the simplest control strategy, if the “ACF only” method is selected at step  820 , then at step  825  the flap actuation frequency and amplitude is set based on the transfer functions obtained from experiment/flight tests. Control voltage U ACF  applied to the piezoelectric actuators will determine the amplitude of flap deflection. This value can be linked to the vibration frequency (f VIB ) and amplitude (P VIB ) and should be set between 0 V and 150 V for the particular design of ACF  106  (see  FIG. 1A ) presented herein. The frequency of actuation can then be linked solely to the frequency of vibration and it should be between (N−1)/rev to (N+1)/rev frequency for best results. APL  100  (See  FIG. 1A ) is idle in this case, with the solid link mode being functional. 
     If the “APL only” method is selected at step  830 , then at step  835  the APL actuation frequency and amplitude will be set based on the transfer functions obtained from experiment/flight tests. Control voltage U APL , however, will not be linked this time to the amplitude of actuation, but to the torsional frequency of rotor blade  104  (see  FIG. 1A ). As has been shown in experiment (see  FIG. 4H ), the resultant stiffness of APL  100  can be set to any value between k 1  and k 2  by setting U APL  between 60 V and 120 V. When an intermediate value is set, APL  100  is in the transitional mode and it extracts energy from the system via sliding friction. This mode of operation is called the “energy extraction mode” and APL  100  is most efficient in this mode when applied on its own (without any Flow Control device) Since the stiffness of APL  100  is linked to the blade resonance frequency in torsion, the blade frequency can essentially be set to any desired value by activating APL  100 . The choice of the desired blade torsional frequency, and thus of U APL , will be driven by the frequency of vibration (f VIB ). The frequency of actuation will also be driven by the frequency of vibration (f VIB ). Note that in this case ACF  106  is idle. 
     Finally, if the hybrid control method is selected at step  840 , both ACF  106  and APL  100  are operational at the same time. First, the ACF operational mode is selected at step  850  based on pilot input or a database, in which vibration levels have been linked to the choice of operational mode. When the Servo Flap mode is selected, then first the ACF actuation parameters (frequency f ACF  and amplitude U ACF ) are determined at step  860  from the transfer functions from experiment, similar to the “ACF only” mode described above. Next, at step  865 , phase angle φ is determined based on experience from tests. The phase angle determines the delay between forcing and response. It is known to be 90 deg at the resonance frequency, whereas it decreases to 0 deg below the resonance frequency and increases to 180 deg above the resonance frequency. The phase angle will dictate that when (in terms of rotor azimuth angle) APL  100  should be activated relative to the actuation of ACF  106  already determined at step  860 . Once the phase angle is known, the APL frequency (f APL ) and control voltage (U APL ) can be determined. Note that in contrast to the “APL only” configuration, these two parameters depend not only on the vibration frequency (f VIB ) but also on phase angle (φ) and the ACF frequency (f ACF ) as well, as shown in step  880 . The method of determining the control parameters for the High-Lift device mode, i.e. steps  870  and  875 , is analogous to the above description, with the difference that f ACF , U ACF  and φ are determined from the transfer function for the high-lift flap mode. 
     The outputs from control computer  710  are the actuation parameters for the Flow Control (i.e. ACF  106 ) and Structural Control (i.e. APL  100 ) systems: these are sent to the control systems at step  890 . Note that the feedback loop between the outputs and the inputs is realized outside of control computer  710  as shown in  FIG. 7 . 
     Thus, in selected embodiments, the Active Pitch Link may serve as a backup system for a “swashplateless” helicopter rotor controlled primarily by a Flow Control device (such as either an Actively Controlled Flap or Active Twist Rotor). Combining such Flow Control device with the Active Pitch Link can have at least two advantages:
         a) the Active Pitch Link can improve the efficiency of the Flow Control device by lowering the torsional stiffness of the blade   b) the Active Pitch Link can serve as a control system backup for the case that the Flow Control device fails. When the Flow Control device fails and is unable to serve its purpose as the primary means of rotor control, the blades (pitch angle) can still be controlled via the Active Pitch Link.       

     Although the hybrid control device of the present invention has been described in relation to rotor blades on a helicopter, it will be understood by those in the art that the invention may be applied to other devices employing blades in which vibration control is desired. For example, the hybrid control device may be applied to the blades of a wind turbine which behaves in a manner similar to a rotor blade such that vibration control would be beneficial.

Technology Classification (CPC): 1