Abstract:
An actuation system for pivoting a flap on a helicopter rotor blade to reduce the interaction of the blade with the preceding blade vortex. The actuation system includes a fluid supply which is connected to first and second fluid supply lines. The fluid supply lines convey flows of pressurized fluid from the fluid supply to an actuator. The actuator includes a housing mounted within the rotor blade and having a channel formed in it. A butterfly shaft is pivotally mounted within the channel and has laterally extending arms which separate the channel into four lobes. A first port connects the first fluid supply line with two diametrically opposed lobes in the channel. A second port connects the second fluid supply line with the other two diametrically opposed lobes in the channel. A torque coupling is attached to the butterfly shaft and engaged with the flap such that rotation of the torque coupling produces concomitant rotation of the flap. The pressurization of the first fluid supply line causes the torque coupling to rotate in a first direction. The pressurization of the second supply fluid line causes the torque coupling to rotate in the opposite direction.

Description:
FIELD OF THE INVENTION 
     The present invention relates to an actuation system for an aircraft and, more particularly, to an improved actively controlled actuator for controlling the flap angle in a helicopter rotor blade. 
     BACKGROUND OF THE INVENTION 
     Helicopter main rotor lift and rotor driving torque produce reaction forces and moments on the helicopter main gearbox. In addition to the primary flight loads, the aircraft is also subjected to vibratory loads originating from the main rotor system. These vibratory loads produce vibrations and noise within the aircraft that are extremely discomforting and fatiguing to the passengers. 
     One vibratory load that is of particular concern results from the interaction of the rotor blades with blade vortices developed by the preceding blades during rotation. As the rotor blade rotates, the air flows passing over and under the blade combine downstream from the trailing edge creating a vortex. During normal flight modes, the blade vortices do not cause any particular problem. However, in certain instances, such as when the aircraft is descending, the trailing blade contacts the blade vortex generating an impulsive noise or slap. This contact with the vortex also creates a vibration within the rotor system that transfers into the cabin. These vibrations can be upwards of 5/rev (i.e., 5 times per revolution of the rotor system). The noise and vibrations generated by the blade interaction with the vortices results in passenger and crew discomfort. 
     Blade vortex interaction also generates an external noise signature which can be easily detected at long range, increasing the aircraft&#39;s vulnerability when in a hostile environment. With the increasing use of helicopters for night reconnaissance missions, it is desirable to minimize the external noise signature of the aircraft. 
     Many attempts have been made over the years to alleviate or reduce blade vortex interactions. A considerable amount of those attempts have been directed toward passive type systems wherein the blade is designed to weaken the vortex at the blade tip. See, for example, U.S. Pat. No. 4,324,530 which discloses a rotor blade with an anhedral swept tapered tip which reduces the intensity and shifts the location of the tip trailing edge vortex so as to reduce the occurrence of blade vortex interactions. 
     While passive solutions have provided some reduction in blade vortex interaction, those solutions also tend to negatively impact the flight characteristics of the rotor blade. 
     Active rotor control systems have recently been proposed to counteract blade vortex interactions. These systems are typically designed to change the motion of the rotor blade to miss the blade vortex or cut the vortex differently so as to reduce contact with the blade vortex. One of these systems is called higher harmonic blade pitch control wherein the blade pitch is controlled to reduce the vortex at the blade tip. While the reduced blade tip vortex does lead to lower noise from blade vortex interaction, the change in blade pitch also reduces the aerodynamic characteristics for the entire blade. 
     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. An actuator is used to control the movement of the flaps and can be either mechanical, electrical, pneumatic, or hydraulic. U.S. Pat. No. 5,639,215 discloses a similar actuatable flap assembly. In this assembly, the actuator is a mechanical actuator that is either a push-rod type device, a linkage, or a servo-motor driven rack. 
     Although the prior art systems for actively controlling the rotor blade interactions with the blade vortex are empirically better than the passive systems described above, these prior art systems do not address the realistic problems associated with mounting an actuation system within a rotor blade to control the flaps in the desired manner. 
     A need, therefore, exists for an improved actuation system for use in an active rotor control system to control flaps on a rotor blade for improving the blade&#39;s aerodynamic performance while reducing noise generating by blade vortex interactions. 
     SUMMARY OF THE INVENTION 
     The present invention relates to an actuator for actuating a flap mounted on a trailing edge of a helicopter rotor blade. The actuator is adapted to be connected to a first and second fluid supply lines. The fluid supply lines provide first and second flows of pressurized fluid from a fluid supply. The actuator includes a housing which is adapted to be mounted within the rotor blade. The housing has a channel formed within it. A butterfly shaft is pivotally mounted within the channel. The butterfly shaft has laterally extending arms which separate the channel into four lobes. 
     A first port formed within the housing is adapted to receive a flow of fluid from the first fluid supply line. The first port is in fluid communication with two diametrically opposed lobes in the channel. 
     A second port formed within the housing is adapted to receive a flow of fluid from the second fluid supply line. The second port is in fluid communication with the other two diametrically opposed lobes in the channel. 
     A torque coupling is preferably attached to the butterfly shaft and adapted to engage with a flap on the rotor blade such that rotation of the torque coupling produces concomitant rotation of the flap. The torque coupling rotates in a first direction when the first port receives pressurized fluid, and rotates in the opposite direction when the second port receives pressurized fluid. 
     An actuation system that includes the above described actuator is also disclosed for actuating a flap on a helicopter rotor blade. 
     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 
     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. 
     FIG. 1 is a plan view of a helicopter rotor blade incorporating an actuatable flap according to the present invention. 
     FIG. 2 is a top schematic view of a portion of a rotor blade illustrating one embodiment of the flap actuation system according to the present invention. 
     FIG. 3 is an enlarged view of one embodiment of a butterfly actuator used to actuate a flap in one embodiment of the invention. 
     FIG. 4A is a cross-sectional view of a helicopter rotor blade incorporating a butterfly actuator for actuating a flap according to the present invention. 
     FIG. 4B is a cross-sectional view of the helicopter rotor blade in FIG. 4A illustrating the flap in an upwardly deflected position. 
     FIG. 4C is a cross-sectional view of the helicopter rotor blade in FIG. 4A illustrating the flap in a downwardly deflected position. 
     FIG. 5 illustrates one embodiment of the butterfly actuator in more detail. 
     FIG. 6 is an isometric view of a test apparatus used to test a flap actuation system according to the present invention, 
     FIG. 7 is an isometric view of an alternate fluid supply system that can be used to supply pressurized fluid to the butterfly actuator. 
     FIG. 8 is a graphical representation of test results showing the hinge moment on a full scale actuator according to the present invention. 
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     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. 
     Certain terminology is used herein for convenience only and is not be taken as a limitation on the invention. Particularly, 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. 
     Referring now to the drawings, wherein like reference numerals illustrate corresponding or similar elements throughout the several views, FIG. 1 illustrates a rotor blade  10  for a helicopter. The rotor blade  10  includes a leading edge  12  and a trailing edge  14 . The blade  10  is attached at its root end  15  to a helicopter hub H and rotated in conjunction therewith about a rotational axis. The blade  10  includes a tip end  16  which is located at the radially outermost point on the blade. The tip end  16  may include a tip cap  17 . At least one flap  18  is mounted to the trailing edge  14  of the blade  10  so as to be articulatable with respect to the trailing edge  14 . As shown, the flap  18  may be located within a recess in the trailing edge  14 . Alternatively, the flap  18  may be located aft of the trailing edge  14  and extent all or partway along the length of the trailing edge  14 . Moreover, there may be several flaps  18  mounted to the rotor blade  10  that can be articulated either independently or concurrently for controlling blade vortex interactions. 
     Referring now to FIG. 2, a schematic plan view of the rotor blade  10  is shown illustrating the actuation system  20  according to one embodiment of the present invention. As discussed above, there have recently been several attempts made at designing an actuation system for controlling flaps on rotor blades. The present invention requires an actuation system that can operate with any rotor system operating at about 2/rev to about 5/rev. The actuation system should be capable of providing flap angular control of about ±10° at approximately 30 cycles/second (Hz) at 5/rev. Furthermore, the actuation system must be capable of providing a sufficient amount of force to overcome the air loads acting on the flap during normal flight. For example, in order to actuate a 69 inch long flap, a torque of 864 in-lbs must be generated to overcome the air loads on the blade. 
     To meet these design requirements, the present invention incorporates an actuation system  20  that includes a fluid actuator. Other types of systems, such as those disclosed in the prior art, were considered but were determined not to be sufficient for use in a full scale rotor system. For example, although an electromechanical actuator (such as a voice coil) or a solid state actuator (such as a piezoelectric) could be used to control flap motion, currently available devices cannot meet the full scale rotor requirements. These requirements include providing a sufficient amount of deflection (e.g., ±10°) and torque, while minimizing weight and staying within the blade airfoil contour to avoid creating an aerodynamic drag penalty. 
     The actuation system  20  includes a fluid supply (not shown in FIG.  2 ), which can include, for example, a pump and a fluid source. While a pneumatic system could be used, it is not preferred since air is compressible. Also, a pneumatic system capable of producing the 2000-3000 psi would have serious safety and operational issues. Hydraulic fluid is more preferred since its compressibility is less than air and, thus, provides better control over the actuator motion. The fluid supply is, in one embodiment, located externally from the rotor blade  10 . Preferably a single fluid supply provides a pressurized fluid medium to the actuation systems  20  located in all the blades  10 . The fluid supply is preferably located within the rotor hub H and rotates in conjunction with the rotor blades  10 . The fluid supply can be connected through an electrical control line to a power source and/or a controller for controlling the fluid supply. The power source and/or controller may be located within the rotor hub or can be located within the aircraft with the power and control commands being transferred from the aircraft to the rotating hub through any conventional means, such as a slip ring connection. 
     The fluid supply  22  is fluidly connected to at least one supply line  26 . The supply line  26  provides pressurized fluid for actuating the flap  18 . In a more preferred embodiment, there are two fluid supply lines  26   A ,  26   B , one supply line provides fluid for actuating the flap upward and the second supply line provides fluid for actuating the flap downward. The fluid supply lines  26   A ,  26   B  extend outward through the rotor blade from the root end  15  toward the tip end  16 . The fluid supply lines  26   A ,  26   B  function as conduits for transferring the fluid medium from the fluid supply to an actuator  28 . The fluid supply lines  26   A ,  26   B  are preferably made from high strength steel with a thin wall to reduce the weight of the supply lines. The supply lines  26   A ,  26   B  must still be sufficiently strong to accommodate the internal pressure caused by the fluid medium. During flight, rotation of the rotor head produces a very high centrifugal pressure on the fluid in the supply lines  26   A ,  26   B . This results in an internal pressure that can reach upwards of several thousand psi and higher. In addition to this high centrifugal pressure is the fluid pressure that the system  20  must supply to actuate the flaps. 
     The actuator  28  is shown mounted within the rotor blade adjacent to the flap  18 . More particularly, in the illustrated embodiment, the actuator  28  is attached to the flap  18  at a flap hinge or pivot axis. Referring to FIG. 3, a cross-sectional view of the actuator is shown in more detail. The actuator  28  is a butterfly actuator that includes a housing  30  with a channel  32  formed within it. A butterfly shaft  34  is pivotally mounted within the channel  32  and includes two arms  37  that separate the channel  34  into four lobes  36   A ,  36   B ,  36   C ,  36   D . Pivoting of the butterfly flap  34  changes the size of each lobe. 
     The actuator  28  also includes at least two sets of ports  38 ,  40  that fluidly communicate with the channel to supply fluid from the supply lines to the lobes. More particularly, the first set of ports  38  permit fluid to flow from the first fluid supply line  26   A  through a manifold to lobes  36   A  and  36   D . The second set of ports  40  permit fluid to flow from the second fluid supply line  26   B  through a manifold to lobes  36   B  and  36   C . 
     The butterfly shaft  34  is attached to a torque coupling  42  (shown in FIG. 2) which, in turn, attaches to the flap  18 . The flap  18  is hinged to the rotor blade  10  so that the flap  18  can be articulated through the desired angular range (e.g., ±10°). Any conventional hinge can be used to mount the flap  18  to the rotor blade  10  which allows for pivotal movement of the flap  18 . Those skilled in the art would readily appreciate the diverse hinge mechanisms that can be used in the present invention. 
     Referring now to FIGS. 4A through 4C, the operation of the butterfly flap is illustrated. In FIG. 4A, the flap is shown in a non-deflected position. In this position, the butterfly actuator  28  is receiving pressurized hydraulic fluid along both fluid supply lines  26   A ,  26   B . The supply lines  26 A,  26 B provide an equal amount of pressurized fluid to each lobe  36   A ,  36   B ,  36   C ,  36   D , resulting in equal and opposite loads being applied to the butterfly arms  37 . As a consequence, no torque is applied to the torque coupling  42 . 
     When it is desired to actuate the flap upwards, the first fluid supply line  26   A  is pressurized and the second fluid supply line  26   B  is depressurized. The pressurized fluid from the first fluid supply line  26   A  fills lobes  36   A  and  36   D , increasing the pressure on the arms  37  of the butterfly shaft  34 . The increased pressure causes the butterfly shaft  34  to pivot (FIG.  4 B), thus producing torque on the torque coupling  42 . The torque coupling  42 , in turn, pivots the flap upward as shown in FIG.  4 B. At the same time that lobes  36   A  and  36   D  are pressurized, lobes  36   B  and  36   C  are depressurized, allowing fluid to flow out of the actuator  28  toward the second fluid supply line  26   B . 
     Similarly, when lobes  36   B  and  36   C  are supplied with pressurized fluid from the second supply line  26   B , the butterfly shaft  34  pivots counter-clockwise as shown in FIG. 4C, thus producing torque on the torque coupling  42 . The torque coupling  42 , in turn, pivots the flap downward. At the same time that lobes  36   B  and  36   C  are pressurized, lobes  36   A  and  36   D  are depressurized, allowing fluid to flow out of the actuator  28  toward the first fluid supply line  26   A . 
     It should be noted that if there are several actuators on a supply line connected to a single flap, then the actuators on each supply line are preferably actuated at the same time. If, however, more than one flap is used in the present invention, than pressurized fluid would be sent to the actuators that control the flap that is to be actuated. 
     FIG. 5 is a top sectional view of butterfly actuator  28  configuration which was built and tested. The butterfly shaft  34  is shown positioned within the housing  30  and attached to the torque coupling  42 . The butterfly shaft  34  is mounted within bearings  44  to permit low friction rotation. The fluid supply lines  26   A ,  26   B  are attached to the manifolds within the housing and supply pressurized fluid to the lobes. End caps  46  are mounted to either end of the housing  30  and seal the butterfly shaft  34  in the housing  30 . 
     The actuator  28  shown in FIG. 5 was incorporated into a test apparatus illustrated in FIG. 6 to test the torque capabilities of the butterfly actuator  28  and the aerodynamic effectiveness of the active flap  18 . The apparatus includes a blade segment  10  with a flap  18  attached to it through a pivotal hinge  50 . The hinge  50  is attached to the torque coupling  42  which, in turn, is attached to the butterfly actuator  28 . The fluid supply lines  26  supply fluid from a servovalve  52 . 
     FIG. 8 shows experimental measurements of the flap hinge moment (equal to actuator torque for 5 degree amplitude oscillations at opening frequencies from 1 per rev (5 Hz) to 6 per rev (30 Hz) on the test apparatus of FIGS. 5 and 6, tested at Mach numbers from 0.4 to 0.75 (560 mph). The results are displayed in terms of a full scale system with a 72 inch span flap on 24 inch cord blade. 
     An alternative embodiment of the flap actuation system is shown in FIG.  7 . In this embodiment, the blade actuation system includes a modified fluid supply system  300 . The system  300  includes a servo control valve  302  which supplies pressurized fluid to a butterfly actuator  304  as described above. The servo control valve  302  can be any conventional valve, such as a servo valve made by Moog, Inc or HR Textron, Inc. Fluid is contained within a fluid accumulator tube  306  that extends radially outward from the root to the tip. The fluid accumulator tube  306  is fluidly connected to the valve  302  and to a pump  308 . The pump is also fluidly connected to the valve  302 . The pump  308  supplies fluid from the valve  302  to the accumulator tube  306 . 
     During flight, the centrifugal loads on the rotor blade cause the fluid within the accumulator to increase in pressure. Pressures of upwards of 400 psi can readily be achieved. Even higher pressures can be achieved if the fluid is pressurized before flight. The pressurized fluid is supplied to the valve  302  for use by the butterfly actuator  304  to pivot the flap  18  as described above. The pump draws hydraulic fluid out of the butterfly actuator  304  and channels it to the accumulator tube  306 . 
     This alternate embodiment of the invention takes advantage of the natural centrifugal loads that exist. The entire fluid supply system  300  would be located within the rotor blade  10 . The system eliminates the need for a fluid interface between the blade and the hub. Only electrical power needs to be supplied to operate the system. The system also requires less hydraulic fluid. However, the location of the pump with respect to the feathering axis and its axial location within the blade can create performance and weight problems that must be considered when mounting the pump. 
     The actuation systems described above provides novel means for actuating a flap in an active control system for a rotor blade. The butterfly configuration permits that actuator to have a low profile so that it fits within the confines of the rotor blade. The butterfly configuration also provides high torque to accommodate the high air loads. 
     Also, the low moment of inertia of the actuator configuration permits quick responses. The shape of the channel and the butterfly shaft provides structural stops which limit the rotation of the shaft. As such, external stops are not needed. Furthermore, the internal surfaces can be coated with a friction reducing coating to improve the actuator performance and reduce wear. 
     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.