Abstract:
A plasma actuator system and method especially well adapted for use on airborne mobile platforms, such as aircraft, for directional and/or attitude control. The system includes at least one plasma actuator having first and second electrodes mounted on a surface of an aircraft. The first and second electrodes are arranged parallel to a boundary layer flow path over the surface. A third electrode is mounted between the first and second electrodes and laterally offset from the first and second electrodes. A high AC voltage signal is applied across the first and third electrodes, which induces a fluid flow between the energized electrodes that helps to delay separation of the boundary layer. Applying the AC voltage across the second and third electrodes causes an induced fluid flow that creates the opposite effect of influencing the boundary layer flow to separate from the surface. A plurality of the actuators can be selectively placed at various locations on the aircraft, and selectively energized to provide directional control and/or attitude control over the aircraft.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     The present application is related in general subject matter to U.S. application Ser. No. 11/753,857, filed May 25, 2007 and U.S. application Ser. No. 11/753,869, filed May 25, 2007. 
     The present disclosure is also related in general subject matter to U.S. application Ser. No. 11/403,252, filed Apr. 12, 2006, and assigned to The Boeing Company. 
     All of the above-mentioned applications are hereby incorporated by reference into the present disclosure. 
     FIELD 
     The present disclosure relates to flow control systems, and more particularly to a plasma flow control system and method that is selectively controllable to help prevent separation of a boundary layer flow from a surface of a mobile platform or object, or to cause separation of the boundary layer flow from the surface. 
     BACKGROUND 
     The statements in this section merely provide background information related to the present disclosure and may not constitute prior art. 
     To be aerodynamically efficient, airborne mobile platforms such as aircraft and weapons (air vehicles) typically must have highly integrated configurations. These configurations typically need to combine good performance and useful payload with good stability and control characteristics. To achieve this objective, air vehicle configurations should have efficient, effective and robust control effector suites. Removing conventional control surfaces to make the air vehicle aerodynamically more efficient provides a unique challenge in air vehicle stability and control. 
     Previous work with air vehicles that are tailless and/or hingeless has proven especially challenging in providing vehicle control, especially directional control of the vehicle. A particular problem with hingeless or tailless control is generating directional control at low to moderate angles of attack, with such angles typically being in the range between about 0-4 degrees. At the present time, most aerodynamic methods used for generating directional control at low to moderate angles of attack on an air vehicle involve the use of vertical tails or deflecting a control surface. Providing directional control at low to moderate angles of attack, if any, is a limitation of prior solutions when the vertical tail is removed. 
     Weight is also an important consideration on many forms of mobile platforms, and particularly airborne mobile platforms such as aircraft. Present day aerodynamic control systems typically employ hinged panels that are deflected to alter the boundary layer flow over a surface of the mobile platform, such as over a trailing edge of a wing. As will be appreciated, hinges and the related linkage and hydraulic or electromechanical actuators needed to employ them can add significant weight to an aircraft, thereby increasing the fuel required for a given flight or mission, or reducing the overall payload of the aircraft. 
     SUMMARY 
     The present disclosure relates to a plasma actuator system and method for use on mobile platforms, and particularly on high speed airborne mobile platforms such as jet aircraft. The plasma actuator system forms a flow control apparatus that is useful for controlling a boundary layer flow over a surface of the mobile platform. 
     In one implementation a method is provided for controlling flight of a mobile platform. The method involves disposing a plasma actuator on a surface of the mobile platform so as to be in a path of a boundary layer flow over the surface. The plasma actuator is controlled to assume a first operating configuration in which the plasma actuator influences the boundary layer flow in a manner to draw the boundary layer toward the surface and maintain the boundary layer flow against the surface. The actuator may also be controlled to assume a second operating configuration in which the plasma actuator influences the boundary layer flow in a manner to cause separation of the boundary layer flow from the surface. 
     In one specific implementation, disposing the plasma actuator involves disposing a plasma actuator having first and third electrodes spaced apart along a direction of flow of the boundary layer. A third electrode is disposed intermediate the first and second electrodes, and within a plane that is laterally offset from a plane in which the first and second electrodes are disposed. A dielectric material is disposed between the third electrode and the first and second electrodes. 
     In one embodiment a system for controlling flight of an airborne mobile platform is disclosed. The system includes a plasma actuator disposed adjacent a surface of a mobile platform and an AC voltage source for electrically energizing the plasma actuator. The plasma actuator has a first electrode disposed adjacent the surface of the mobile platform so as to be in a path of a boundary layer flow over the surface, and a second electrode disposed adjacent the surface downstream of the first electrode, relative to a direction of flow of the boundary layer. A third electrode is separated from the first and second electrodes by a dielectric layer, and is disposed between the first and second electrodes and within a plane that is laterally offset from the first and second electrodes. A controller controls the application of an AC voltage from the AC voltage source to the electrodes to at least one of:
         apply the AC voltage across the first and third electrodes, to cause ionization of air between the first and third electrodes that delays separation of the boundary layer flow on the surface; and   apply the AC voltage across the second and third electrodes, to cause ionization of air between the second and third electrodes that causes separation of the boundary layer flow on the surface.       

     In one embodiment the system and method forms a plasma actuator that is able to selectively prevent separation of a boundary layer flow from a surface of an object, as well as to cause separation of the boundary layer flow. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The drawings described herein are for illustration purposes only and are not intended to limit the scope of the present disclosure in any way. 
         FIG. 1  is a plan view of a mobile platform incorporating a plurality of plasma actuators in accordance with one embodiment of the present disclosure, where the plasma actuators are employed along the leading edges of the wings of an aircraft; 
         FIG. 2  is an enlarged cross-sectional side view of one of the plasma actuators shown in  FIG. 1  taken in accordance with section line  2 - 2  in  FIG. 1 , illustrating the actuator energized to delay separation of the boundary layer flow on the surface of the wing, and also showing in simplified form the AC voltage source and the controller used to control the actuator; and 
         FIG. 3  is view of the plasma actuator of  FIG. 2 , but with the actuator being controlled to cause separation of the boundary layer flow from the surface of the wing. 
     
    
    
     DETAILED DESCRIPTION 
     The following description is merely exemplary in nature and is not intended to limit the present disclosure, application, or uses. 
     Referring to  FIG. 1 , there is shown a mobile platform, in this example an aircraft  12 , incorporating a plurality of plasma actuators  10 . In this example the plasma actuators  10  are disposed adjacent leading edges  16   a  and  16   b  of the wings  14   a  and  14   b , respectively, of the aircraft  12 . It will be appreciated, however, that the plasma actuators  10  may be used on virtually any form of mobile platform where it is desirable to effect directional or attitude control of the mobile platform without the need for hinged or moveable panels. Other possible applications may involve unmanned aircraft, missiles, rotorcraft, high speed land vehicles, and possibly even high speed marine vessels. Also, while the plasma actuators  10  are illustrated on the wings  14   a , 14   b  of the aircraft  12 , they could just as readily be employed along the fuselage, on the horizontal stabilizers, the vertical tail, boattail or any other location where it may be desirable to influence the boundary layer flow over the aircraft. 
     In practice, it will often be desirable to include a plurality of the plasma actuators  10  along a surface where control over the boundary layer is needed, as illustrated in  FIG. 1 . The spacing between adjacent plasma actuators  10 , the dimensions of the actuator, and the specific number of actuators, will be determined by the needs of a specific application. 
     Referring to  FIG. 2 , one of the plasma actuators  10  is shown from a side cross-sectional view. Each plasma actuator  10  includes a first electrode  18 , a second electrode  20  and a third electrode  22 . The second electrode  20  is spaced apart from the first electrode  18 . Preferably, the electrodes  18  and  20  are recessed mounted in a surface  24  of the wing  14   a  so that upper surfaces  18   a  and  20   a  of the electrodes  18  and  20 , respectively, are positioned generally flush with the surface  24 , and co-planar with one another. Alternatively, the electrodes  18  and  20  could be mounted on top of the surface  24 . Recess mounting of the electrodes  18  and  20 , however, will help to maintain the original aerodynamic profile of the wing  14   a  or other surface with which the actuator  10  is being implemented on and to reduce aerodynamic drag. 
     The third electrode  22  is mounted between the first electrode  18  and the second electrode  20 , and generally longitudinally in line with the electrodes  18  and  20 , but is disposed so that it sits laterally offset from (i.e., elevationally below) the electrodes  18  and  20 . A dielectric layer of material  25  is provided around the third electrode  22  that separates it from the first and second electrodes  18  and  20 , respectively. Each of the electrodes  18 ,  20  and  22  may be formed with a generally rectangular shape having its major (i.e., long side) axis arranged perpendicular to the direction of flow of the boundary layer. Other orientations are possible as well, depending on the needs of a specific application. 
     In practice, the electrodes  18 , 20 , 22  may be formed from any conductive material. Copper is one material that is particularly suitable. The electrodes  18 , 20 , 22  may be formed as thin strips, possibly as foil strips, and may have a typical thickness on the order of about 0.001-0.005 inch (0.0254-0.127 mm). The length and width of each electrode  18 , 20 , 22  may vary as needed to suit specific applications, but it is anticipated that in many aircraft applications, the length and width of each electrode may typically be on the order of 1-20 inches (2.54 cm-50.08 cm) for the length and 0.12-0.20 inch (3-5 mm) for the width for each of the electrodes  18  and  20 . The width of the buried electrode  22  will typically be wider than that employed for the electrode  22 , and typically on the order of 1.0-2.0 inches (2.54 cm-5.08 cm) depending on the operating voltage being supplied by the AC voltage source  26 . The dielectric layer of material  25  may comprise any suitable dielectric material, for example quartz, KAPTON® or TEFLON® dielectric materials. Other dielectric materials such as ceramics may also be suitable for use, and the precise dielectric used may be dictated by the needs of a specific application. A portion of the dielectric layer of material  25  may also be used to fill the gap between the first and second electrodes  18  and  20 . The elevational spacing of the third electrode  22  from the first and second electrodes  18  and  20 , will typically be about 0.003-0.50 inch (0.076-12.7 mm), although this may also vary significantly as well depending on the needs of a specific application. 
     With further reference to  FIG. 2 , an AC voltage source  26  is coupled to the third electrode  22  and through a pair of switches  28  and  30  to the first and second electrodes  18  and  22 , respectively. The AC voltage source  26  generates a low current, high voltage AC signal, preferably in the range of about 3,000-20,000 volts. The frequency of the AC voltage source  26  is typically between about 1 KHz-20 KHz, but may vary as needed to meet a specific application. The precise output from the AC voltage source  26  is preferably variable to enable the actuator  10  to provide a variable degree of fluid flow control. 
     A controller  32  is in communication with the switches  28  and  30 . The switches  28 , 30  may be semiconductor switching devices suitable for handling the voltage generated by the AC voltage source  26  or may comprise any other suitable forms of switching devices. As will be described in further detail in the following paragraphs, components  18 ,  20 ,  22 ,  25 ,  26 ,  28 ,  30  and  32  effectively form a “dual mode” plasma actuator apparatus that is able to selectively cause or inhibit separation of the boundary layer from the surface  24 . The controller  32  may also be used to control the precise output from the AC voltage source  26 . In one implementation, the controller  32  may be used to control the switches  28  and  30  to generate AC voltage pulses that are applied across the electrode pairs  18 , 22  and  20 , 22 , with a duty cycle between about 10%-100%. Applying a pulsed AC signal to the electrode pairs  18 , 22  and  20 , 22  may result in an increase in power efficiency and overall effectiveness of the actuator  10 . 
     Referring further to  FIGS. 2 and 3 , the operation of the plasma actuator  10  will be described. In  FIG. 2 , when it is desired to prevent separation of the boundary layer flow from the surface  24 , the controller  32  causes switch  28  to be energized (i.e., closed) and switch  30  to be opened. This results in the high AC voltage from the AC voltage source  26  being applied across electrodes  18  and  22 . The high voltage causes air in the vicinity of the spacing between electrodes  18  and  22  to be ionized. Ionization typically occurs when an AC voltage of about 3,000 volts is applied across the electrodes  18  and  22 . The electric field that is created acts on the ionized air to accelerate the charged particles, which collide with the neutral boundary layer air molecules to create a “wall jet”. The strength of the electric field is directly proportional to the magnitude of the applied AC voltage. More particularly, the electric field induces a body force impulse on the ionized air that serves to induce a fluid flow (i.e., the wall jet) very near the surface  24 . The induced fluid flow is indicated by arrow  34 . The induced fluid flow  34  causes an increase in the momentum of the boundary layer fluid near the surface  24 . The resulting induced fluid flow is from the first electrode  18  toward the third electrode  22 . The induced fluid flow  34  functions to prevent, or at least significantly delay, separation of the boundary layer from the surface  24 . Accordingly,  FIG. 2  illustrates what may be viewed as an “attached flow mode” or “first operating configuration” for the actuator  10 . 
     Referring to  FIG. 3 , when it is desired to cause separation of the boundary layer from the surface  24 , the controller  32  energizes (i.e., closes) switch  30  and opens switch  28 . This also causes the air in the region between the second electrode  20  and the third electrode  22  to ionize, but the induced fluid flow, represented by arrow  36 , is in a direction generally opposite to the induced fluid flow  34 . The induced fluid flow  36  serves to cause separation of the boundary layer from the surface  24 . Thus, simply by controlling which pair of electrodes  18 , 22  or  20 , 22  of each plasma actuator  10  the AC voltage is applied across, the boundary layer flow can be influenced as needed. When a variable AC voltage is applied, then the strength of the electric field, and thus the degree to which the electrodes  18  and  22  influence the boundary layer flow, can be varied.  FIG. 3  illustrates what may be viewed as a “separated flow mode” or “second operating configuration” for the actuator  10 . 
     The plasma actuators  10  may be used for directional control purposes, for instance at low angle of attack, by controlling the actuators  10  on the wings  14   a  and  14   b  differently. For example, by controlling the plasma actuators  10  on wing  14   a  so that one effect is achieved, for example preventing flow separation, while controlling the actuators  10  on wing  14   b  to induce flow separation, directional control of the aircraft  12  can be achieved. The directional control results from the differential drag produced by the cooperative effects of the plasma actuators  10  on the wings  14   a  and  14   b , and the moment arm generated at each wingtip about the centerline of each wing  14   a  and  14   b.    
     As should be apparent, the above is merely one example of how the plasma actuators  10  may be implemented on the aircraft  12 . The plasma actuators  10  may instead be used to generate a differential side force on the fuselage of an aircraft or missile, and thus generate a yawing moment. Alternatively, a differential lift could be generated at the wings  14   a  and  14   b  to induce a roll moment. 
     The elimination or reduction of conventional mechanical/hydraulic drive control effectors can significantly reduce the weight of an aircraft, and thus produce increased mission flight time or range for a given aircraft. The plasma actuators  10  and related system and methodology described herein may be used to replace conventional control effectors such as leading or trailing edge flaps, ailerons, moving tail surfaces and vortex generators, thus reducing weight and drag associated with such components. 
     While various embodiments have been described, those skilled in the art will recognize modifications or variations which might be made without departing from the present disclosure. The examples illustrate the various embodiments and are not intended to limit the present disclosure. Therefore, the description and claims should be interpreted liberally with only such limitation as is necessary in view of the pertinent prior art.