Abstract:
Systems and methods for controlling air vehicle boundary layer airflow are disclosed. Representative methods can include applying electrical energy bursts and/or other energy bursts in nanosecond pulses in the boundary layer along a surface of an air vehicle. In a particular embodiment, electrical energy is discharged into the boundary layer to reduce the tendency for the boundary layer to separate and/or to reduce the tendency for the boundary layer to transition from laminar flow to turbulent flow. In other embodiments, energy can be discharged via pulses having a pulse width of about 100 nanoseconds or less, and an amplitude of about 10,000 volts or more. Actuators discharging the energy can be arranged in a two-dimensional ray of individually addressable actuators. Energy can be delivered to the boundary layer via a laser emitter, and energy can be received in a receiver after having transited over at least a portion of the airflow surface. In another embodiment, high energy electrons can be injected into the boundary layer using a hollow cathode array at the airflow surface. In still another embodiment, energy can be introduced at the surface of the air vehicle at a rate sufficient to heat the flow and cause shock waves to propagate into the flow.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    The present application claims priority to U.S. Provisional Application No. 61/019,202, filed Jan. 4, 2008 and incorporated herein by reference. 
     
    
     TECHNICAL FIELD 
       [0002]    The present disclosure is directed generally to systems and methods for controlling flows with pulsed discharges, including via dielectric barrier discharge generators, plasma actuators, hollow cathode arrays, particle accelerators, UV light, laser emitters, and/or other devices. 
       BACKGROUND 
       [0003]    During flight, a boundary layer of air builds up on the exposed surfaces of an aircraft. The boundary layer is a thin film of low velocity, low dynamic pressure air located near a solid boundary and resulting from the air being at rest along the solid boundary. The boundary layer which forms on surfaces located upstream of an aircraft engine can become ingested by the engine and decrease the recovery of total pressure and corresponding thrust performance. Further, the ingested boundary layer increases the flow distortion (a measurement of the quality or uniformity of flow characteristics) at the engine and thereby decreases the stability of engine operation. On the aircraft wing and/or other external surfaces of the aircraft, the boundary layer can increase skin friction and therefore drag. In some instances, the boundary layer can cause premature separation of the flow from the external surface, further increasing drag and/or reducing lift. 
         [0004]    As a result of the foregoing drawbacks associated with boundary layers, many aircraft have employed some type of boundary layer removal, reduction, and/or control system to provide for stable engine operation and increased aerodynamic performance. Representative systems include boundary layer diverters, “bump” boundary layer deflectors, boundary layer bypass ducts, vortex generators, and porous surfaces or slots that either bleed boundary layer flow from the surface, or energize the flow by air injection. Unfortunately, these systems are often complex and can entail a substantial increase in aircraft weight and/or volume. 
         [0005]    One recent technique for addressing boundary layer flow is to use a dielectric barrier discharge device to energize and/or redirect the boundary layer flow. These devices operate by ionizing air adjacent to the flow surface in such a way as to generate or direct flow adjacent to the surface. Accordingly, dielectric barrier discharge devices typically include a pair of electrodes separated by a dielectric material. The voltage applied to at least one of the electrodes is typically cycled in a sinusoidal fashion to ionize the adjacent air. While the foregoing approach has been shown to create the desired effect on the boundary layer at relatively low speeds for realistic air vehicle applications, there remains a need for devices that better control boundary layer flow and do so in a manner that is more efficient and effective than techniques associated with existing devices, and that can operate at higher flow speeds associated with realistic aircraft operations. 
       SUMMARY 
       [0006]    The following summary is provided for the benefit of the reader only and is not intended to limit the disclosure in any way. The present disclosure is directed generally to systems and methods for controlling flows with pulsed discharges. A method for controlling air vehicle airflow in accordance with a particular embodiment includes forming a boundary layer on a surface of an air vehicle. The method further includes reducing a tendency for the boundary layer to separate, and/or reducing a tendency for the boundary layer to transition from laminar flow to turbulent flow, by activating different individually addressable actuators in different manners. The individually addressable actuators can be arranged in a two-dimensional array at the surface. In further particular embodiments, activating different individually addressable actuators includes activating different plasma actuators, or different hollow cathode actuators. 
         [0007]    An apparatus for controlling air vehicle airflow in accordance with another embodiment includes an airflow surface and an array of individually addressable actuators positioned at the airflow surface, within individual actuators coupled to a power source to deliver energy into boundary layer flow adjacent to the airflow surface. The apparatus can further include a controller that is operably coupled to the individual actuators and is programmed with instructions for activating different individually addressable actuators in different manners. For example, individual actuators can be activated at different times and in a particular embodiment, the controller can be programmed with instructions to sequentially activate the actuators in a generally streamwise direction. 
         [0008]    Still another embodiment of the disclosure is directed to an apparatus for controlling air vehicle airflow, and includes an airflow surface and a radiation emitter positioned to direct radiation generally parallel to the airflow surface. The apparatus can further include a receiver spaced apart from the emitter to receive radiation after the radiation has transited over at least a portion of the airflow surface. For example, the radiation emitter can include a laser emitter and, in a further particular embodiment, can emit radiation at wavelengths of greater than about 200 nanometers. 
         [0009]    The features, functions, and advantages that have been discussed can be achieved independently in various embodiments of the present invention or may be combined in yet other embodiments, further details of which can be seen with reference to the following description and drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0010]      FIG. 1  is a schematic illustration of a flow control system, including components of an actuator, configured in accordance with an embodiment of the disclosure. 
           [0011]      FIGS. 2A and 2B  illustrate representative wave forms of electrical signals applied to an actuator in accordance with an embodiment of the disclosure. 
           [0012]      FIG. 3  is a table identifying representative aircraft and corresponding locations at which actuators may be used to control flow in accordance with particular embodiments of the disclosure. 
           [0013]      FIG. 4A  illustrates a fixed-wing aircraft identifying locations at which actuators may be positioned in particular embodiments. 
           [0014]      FIG. 4B  illustrates several rotor craft with corresponding locations at which actuator devices may be positioned in particular embodiments. 
           [0015]      FIGS. 5A-5D  are tables illustrating actuator locations for air vehicles in accordance with further particular embodiments. 
           [0016]      FIGS. 6A-6B  are schematic illustrations of a blended wing body aircraft having actuators located at a forebody surface in accordance with a particular embodiment of the disclosure. 
           [0017]      FIG. 7  is a schematic illustration of an aircraft inlet having a flow control system configured in accordance with an embodiment of the disclosure. 
           [0018]      FIG. 8  is a partially schematic cross-sectional illustration of an inlet having a flow control assembly configured in accordance with another embodiment of the disclosure. 
           [0019]      FIG. 9  is a partially schematic, cross sectional side view of an inlet having a flow control assembly configured in accordance with still another embodiment of the disclosure. 
           [0020]      FIGS. 10A-10C  are partially schematic illustrations of airfoils having flow control systems configured in accordance with further embodiments of the disclosure. 
           [0021]      FIGS. 11A-11B  are partially schematic illustrations of actuators arranged in a grid array in accordance with another embodiment of the disclosure. 
           [0022]      FIG. 12  is a partially schematic illustration of an airfoil having laser beam actuators arranged in accordance with a particular embodiment of the disclosure. 
       
    
    
     DETAILED DESCRIPTION 
       [0023]    The following description is directed generally toward systems and methods for controlling flows with pulsed discharges, for example, via plasmas generated by dielectric barrier discharge devices, other plasma actuators, and/or other devices including laser emitters, hollow cathode arrays, and particle accelerators. Several details describing structures or processes that are well-known and often associated with aspects of these systems and methods are not set forth in the following description for purposes of brevity. Moreover, although the following description sets forth several representative embodiments, several other embodiments can have different configurations or different components than those described in this section. As such, other embodiments of the disclosure may have additional elements or may eliminate several of the elements described below with reference to  FIGS. 1-12 . 
         [0024]    Several embodiments of the disclosure described below may take the form of computer-executable instructions, including routines executed by a programmable computer (e.g., a controller). Those skilled in the relevant art will appreciate that one or more embodiments can be practiced on computer systems other than those shown and described below. Instructions can be embodied in a special-purpose computer or data processor that is specifically programmed, configured or constructed to perform one or more of the computer-executable instructions described below. Accordingly, the term “computer” as generally used herein, refers to any data processor, and can include controllers, multi-processor systems, processor-based or programmable consumer electronics, network computers, mini-computers and the like. 
         [0025]      FIG. 1  is a partially schematic, cross-sectional illustration of a flow control assembly  130  that includes a representative actuator  131  configured in accordance with an embodiment of the disclosure. The flow control assembly  130  can include more than one actuator  131 , but a single actuator  131  is shown in  FIG. 1  for purposes of illustration. The actuator  131  can include a first electrode  133 , a second electrode  134 , and a dielectric material  135  positioned between the first and second electrodes  133 ,  134 . Accordingly, the first and second electrodes  133 ,  134  are separated by a gap  136 . 
         [0026]    In a particular embodiment, the first electrode  133  is located upstream (with reference to a local air flow direction F) from the second electrode  134 , and the upper surface of the first electrode  133  is typically flush with the surrounding flow surfaces. The first electrode  133  is also typically “exposed” to the flow. As used in this context, “exposed” means that the first electrode  133  is in direct electrical contact with the flow, or at least more direct electrical communication with the flow than is the second electrode  134 . The exposed first electrode  133  can accordingly include a protective coating or other material that restricts or prevents erosion due to environmental conditions, without unduly impacting the electrical communication between the first electrode  133  and the adjacent flow, e.g., without unduly impacting the ability to provide direct current coupling between the first electrode  133  and the adjacent flow. In other embodiments, the material forming the first electrode  133  can be selected to have both suitable electrical conductivity and suitable resistance to environmental factors. A representative material includes stainless steel. 
         [0027]    In particular embodiments, the first electrode  133  can include a conductive environmental coating. For example, the first electrode  133  can include a coating formed from a thin layer of tungsten, tungsten carbide (or another tungsten alloy), nichrome or stainless steel. In other embodiments, the coating can include a semiconductive material that becomes conductive as the high voltages described above are applied to the first electrode  133 . For example, the first electrode  133  can include a silicon or gallium arsenide bulk material treated with a suitable dopant (e.g., boron or phosphorus, in the case of silicon). In other embodiments, other suitable conductive and/or semiconductive materials can be applied to the first electrode. In any of these embodiments, the material can be selected to provide the necessary level of conductivity and the necessary resistance to environmental conditions, including resistance to rain erosion, oxidation and exposure to fuel and/or ice protection chemicals. 
         [0028]    It is expected that the majority of the electric field lines emanating from the first electrode  133  will emanate from the trailing edge of the electrode. Accordingly, in at least some cases, the environmental coating can be applied to the majority of the exposed surface of the first electrode  133 , leaving only a small, aft portion of the first electrode  133  uncoated. In such cases, the coating may be selected to be entirely non-conductive (e.g., a dielectric coating) without causing undue interference with the ionizing electrical field emanating from the first electrode  133 . 
         [0029]    The second electrode  134  can be covered or at least partially covered with the dielectric material  135 , for example, to prevent direct arcing between the two electrodes. The first electrode  133  or the second electrode  134  is coupled to a controller  137 , which is in turn coupled to a power supply  138  to control the power delivered to the first electrode  133  or the second electrode  134 . The other electrode  133  or  134  may also be coupled to the power supply  138  and/or the controller  137 , or may simply be grounded. The controller  137  can include a computer having a computer-readable medium programmed with instructions to direct a signal waveform to the first electrode  133 , in a manner that is expected to enhance the efficiency and/or the effectiveness of the actuator  131 . 
         [0030]    In many instances, it is expected that relatively high voltage, narrow-width pulses can have a beneficial effect on boundary layers by delaying the transition from laminar to turbulent flow in the boundary layer, and/or by delaying the point at which the boundary layer separates from the surface adjacent to which it flows. Further details of representative pulsed discharge actuators are included in AIAA paper 2007-941, titled “Pulsed Discharge Actuators for Rectangular Wing Separation Control” (Sidorenko et al.) presented at the 45th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nev., Jan. 8-11, 2007, and incorporated herein by reference. Suitable pulse generators are available from Moose Hill Enterprises, Inc. of 54 Jennie Dade Lane, Sperryville, Va. 22740. 
         [0031]      FIG. 2A  illustrates a representative wave form  150  that can be applied by the controller  137  ( FIG. 1 ) to the actuator  131  ( FIG. 1 ). The wave form  150  includes a series of pulse bursts  151 , each of which includes one or more pulses  152 . The pulse bursts  151  are spaced apart in accordance with a modulation frequency  153 . In a particular embodiment, the modulation frequency  153  is selected to match or in another manner correspond with a characteristic fluid instability frequency  154  present in the flow adjacent to the actuator from which the signals emanate. For purposes of illustration, reference numerals  153  and  154  point to periods in  FIG. 2A , but are discussed herein in the context of frequencies. The fluid instability to which the frequencies  153 ,  154  correspond can include Tollmein-Schlichting instabilities, eddy instabilities, sheer layer instabilities, and/or other fluid instabilities. In general, the modulation frequency  153  can be selected to match the characteristic fluid instability frequency  154 , resulting in a Strouhal number of about 1.0. The phase relationship between the pulse bursts  151  and the fluid instabilities can be varied and selected to produce the desired effect on the boundary layer. For example, in some embodiments, the pulse bursts  151  can be timed to be approximately  1800  out of phase with the maximum amplitude of the fluid instabilities. In general, it is expected that selecting the Strouhal number to be approximately 1.0 will produce the greatest impact on the boundary layer, but in other embodiments, the modulation frequency  153  can be selected to have other relationships relative to the characteristic fluid instability frequency  154  and accordingly can have Strouhal numbers other than 1.0. 
         [0032]      FIG. 2B  is an enlarged illustration of one of the pulse bursts  151  shown in  FIG. 2A . For purposes of illustration, the pulse bursts  151  shown in  FIGS. 2A and 2B  have four pulses  152 , but in other embodiments, the pulse bursts  151  may have a greater or lesser number of pulses  152 . Each of the pulses  152  can have a relatively high voltage and short duration, and can be monophasic (e.g., delivered at a single polarity). For example, as shown in  FIG. 2B , each of the pulses  152  can have a pulse duration of from about 10 nanoseconds to about 100 nanoseconds, (e.g., 10-20 nanoseconds) and an amplitude of from about 10 kilovolts to about 60 kilovolts (e.g., 20, 30, 40 or 50 kilovolts). In one embodiment, the pulses  152  can be produced at a frequency of about 6 kHz, but in other embodiments, the frequency of the pulses  152  can be much higher. For example, in representative embodiments, the pulses  152  can have a frequency of from about 10 kHz to about 100 kHz. It is expected that the monophasic pulses will be more likely to have a beneficial effect on the adjacent boundary layer flow than conventional biphasic sinusoidal AC pulses. 
         [0033]    The duty cycle in accordance with which the pulses  152  are produced (e.g., the percentage of time that the pulses are active or “on”) can vary from about 10 percent to about 100 percent. A duty cycle of 100% indicates that the pulses  152  are continually active. In general, it is desirable to have a duty cycle of less than 100% to conserve power, to produce distinct pulses  152 , and to produce pulse bursts  151  that are spaced apart in time in a manner that corresponds to the characteristic fluid instability frequency  154  shown in  FIG. 2A . Accordingly, the duty cycle is generally a function of the pulse width, the frequency at which individual pulses  152  are generated, the number of pulses  152  in the pulse bursts  151 , and the spacing between pulse bursts  151 . These parameters can be selected to produce the desired effect on the boundary layer, while consuming as small an amount of power as is required to have the desired effect. The foregoing characteristics (e.g., the modulation frequency  153 , the number of pulses  152  in a pulse burst  151 , and the pulse width and frequency of the pulses  152 ) can be varied depending upon the conditions of the boundary layer. For example, the wave form  150  may have different characteristics when delivered into a low speed flow than when delivered into a high speed flow. Accordingly, the controller  137  ( FIG. 1 ) can be programmed with instructions that automatically control the delivery parameters in a manner that depends on local flow characteristics, and/or other factors. 
         [0034]    In general, it may be desirable to keep pulse widths as short as possible (while still maintaining flow control) to conserve power and/or to prevent arcing between adjacent electrodes. In some cases, it may be advantageous to deliver pulses having pulse widths up to 100 nanoseconds or above (so long as the electrodes do not arc unacceptably), so as to ease the manufacturability constraints on the discharge devices. 
         [0035]      FIG. 3  is a chart illustrating representative implementations of devices such as the actuator  131  shown in  FIG. 1 . As shown in  FIG. 3 , the actuators can be applied to a number of aerodynamic surfaces and a number of types of air vehicles to produce enhanced results, based generally on increased control over the boundary layer on the aerodynamic surfaces. 
         [0036]      FIG. 4A  is an isometric illustration of a fixed wing aircraft, illustrating representative implementations of actuators in accordance with a variety of embodiments of the disclosure. Such actuators may be positioned at the aft body of the aircraft fuselage to reduce drag, and/or on the trailing edge devices to improve aerodynamic performance of these devices by controlling the boundary layer flowing over these devices. Such actuators may also be placed on the leading edge high lift devices to produce a similar effect. Actuators may be positioned on or near the landing gear to control airflow separation from the landing gear and reduce buffeting. Actuators may also be placed over the surface of the wing to control the boundary layer flow over the wing, and in at least some cases, effectively change the camber of the wing aerodynamically (e.g., without changing the solid contours of the wing), thus allowing the camber of the wing to be adjusted at different flight conditions by selectively activating the actuators and/or selectively changing the manner in which the actuators deliver energy into the adjacent flow. Actuators may also be positioned at the vertical stabilizer to provide additional control of the airflow adjacent to the vertical stabilizer and, in at least some cases, allow the vertical stabilizer to have a reduced size as a result. 
         [0037]      FIG. 4B  illustrates representative actuator installations for rotor craft. The actuators can be installed on the rotor mast to reduce drag, on the vertical stabilizer (for rotor craft that include a vertical stabilizer) to reduce stabilizer size, on rotor pylons to reduce drag on the pylons, on the rotor craft aft body to reduce aft body drag, and/or on the rotors themselves to improve rotor performance. In a particular embodiment, the actuators may be positioned on the rotors to reduce the likelihood for separation during the downwind portion of the rotor rotation cycle, e.g., to prevent or delay retreating blade stall. Such an installation is expected to have a significant beneficial effect on the performance of the rotors by enhancing lift and/or reducing drag, and/or reducing rotor vibration (e.g., by dynamically controlling unsteady pressure loads on the rotor). 
         [0038]      FIGS. 5A and 5B  further illustrate representative installations of actuators in accordance with particular embodiments of the disclosure on fixed wing aircraft. The actuators can have beneficial effects for a wide variety of installations on both military and commercial aircraft.  FIG. 5C  further illustrates representative actuator installation locations for rotor craft.  FIG. 5D  identifies representative actuator installation locations for inlets and associated flow surfaces (e.g., inlet forebodies). 
         [0039]      FIG. 6A  illustrates a representative aerodynamic body  100  on which a flow control assembly  130  is installed in accordance with a particular embodiment. The aerodynamic body  100  can include an aircraft  101  having an airfoil  120  and one or more engines  114  (three are shown in  FIG. 6A  for purposes of illustration). The aircraft  101  shown in  FIG. 6A  has a blended wing-body configuration with the engines  114  located aft. In other embodiments, the aerodynamic body  100  can have other configurations capable of atmospheric flight (such as those shown in  FIGS. 4A-5D ), including, without limitation, tube and wing configurations (typical of commercial transports and private aircraft), missile configurations, or rotorcraft configurations. The aircraft  101  can be manned or unmanned. In a particular embodiment shown in  FIG. 6A , the flow control assembly  130  is positioned to control the flow entering the engines  114 . In other embodiments, the flow control assembly can be positioned to control the flow over other portions of the aircraft, including the airfoil  120 , as will be discussed later. 
         [0040]    Air is supplied to the engines  114  via an air induction system  110 . The air induction system  110  can include one or more inlets  111  (e.g., one inlet  111  per engine  114 ), each having an inlet aperture  112  and an inlet duct  113  that directs air in an aft direction to the engine  114 . In the configuration shown in  FIG. 6A , the inlets  111  are located at a point well aft of a forward leading edge  121  of the airfoil  120 . Accordingly, a surface  102  of the aerodynamic body  100  is positioned upstream of the apertures  112  such that intake air moves over the surface  102  prior to being received by the inlets  111 . A boundary layer of low velocity air builds up on the surface  102  beginning at the leading edge  121 , and moves in a generally aft direction toward the inlets  111 . The illustrated flow control assembly  130  is positioned to control the boundary layer air before it enters the inlets  111 . Optionally, the flow control assembly  130  can be positioned to control the flow within the inlets  111 , in addition to or in lieu of controlling the flow external to the inlets  111 . 
         [0041]    As shown in  FIG. 6B , the inlets  111  can have a generally round shape that may or may not be offset upwardly away from the flow surface  102 . When the inlets  111  are offset away from the flow surface  102 , a diverter can be positioned between the inlets  111  and the flow surface  102  to remove some or all of the boundary layer flow. However, in many cases, it is desirable not to offset the inlets  111  from the flow surface  102  so as to reduce weight and drag. The flow control assembly  130  ( FIG. 6A ) can replace the diverter (or at least reduce the size of the diverter) in the foregoing instances, as described below with reference to  FIG. 7 . 
         [0042]      FIG. 7  is a top perspective illustration of a portion of an integrated, non-axisymmetric, diverter-less inlet  711 . In one aspect of this embodiment, the inlet  711  includes an inlet aperture  712  and an inlet duct  713  positioned aft of the inlet aperture  712 . The inlet aperture  712  has a non-axisymmetric shape and is positioned aft of a forebody surface  715 . Accordingly, a flow surface  702  directing air into the inlet aperture  712  and then aft to the engine (not visible in  FIG. 7 ) can include portions of the forebody  715  and/or portions of the inlet duct  713 . 
         [0043]    A flow control assembly  730  is positioned at the flow surface  702  to control the flow entering the inlet  711 . In a particular aspect of an embodiment shown in  FIG. 7 , the flow control assembly  730  is installed at the forebody  715 . In other embodiments, portions of the flow control assembly  730  may be installed in the inlet duct  713 , in addition to or in lieu of the location at the forebody  715 . The flow control assembly  730  in the illustrated embodiment includes multiple actuators  731  (e.g., dielectric barrier discharge devices, plasma actuators, other electrically operated ionizing devices, and/or other energy-emitting devices) arranged forward of the inlet  711 . Individual actuators  731  may be arranged in rows  732  (two of which are specifically identified as rows  732   a  and  732   b ) that are oriented at least partially transverse to incoming flow streamlines  722  and to the inlet aperture  712 . In still a further particular aspect of the arrangement shown in  FIG. 7 , the rows  732  can be angled relative to the incoming streamlines  722  so as to form a “chevron” pattern. Accordingly, at least some of the incoming boundary layer flow can be directed outboard, away from the inlet  711  (as indicated by flow streamlines  722   a ) and/or some of the flow directed into the inlet  711  can be energized (as indicated by streamline  722   b ). The number of rows  732  of actuators  731  can be selected to be as small as possible while still providing the desired boundary layer diversion and/or energizing effect. For example, the forebody  715  can include five rows  732  of actuators  731  positioned forward of the inlet  711 . In other embodiments, the number of rows  732  can be different, depending upon the specific geometry into which the actuators  731  are integrated. 
         [0044]      FIGS. 8 and 9  illustrate inlets having flow control assemblies configured in accordance with other embodiments of the invention. For example,  FIG. 8  illustrates a generally axisymmetric inlet  811  (such as is typically used on a commercial jet aircraft) having a generally circular inlet aperture  812  and a generally axisymmetric inlet duct  813  positioned forward of an engine  814 . An associated flow control assembly  830  includes actuators  831  positioned in multiple circumferential rows  832  between the inlet aperture  812  and the engine  814 . For purposes of illustration, only the top-most and bottom-most actuators  831  are shown for each row  832 . Each actuator  831  can have a configuration generally similar to that described above with reference to  FIG. 1 , and can have applied to it a waveform generally similar to that described above with reference to  FIGS. 2A-2B . The actuators  831  can be arranged in a series of five rows, as shown in  FIG. 8 , or fewer rows. In any of these embodiments, the effect of the actuators  831  is to control the flow entering the inlet  811  so as to reduce the likelihood for flow separation, increase the total pressure recovery, and/or reduce flow distortion at the engine entrance. 
         [0045]      FIG. 9  illustrates another representative inlet  911  having a forebody  915  positioned forward of an inlet aperture  912 , and an adjacent or aft surface  916  positioned aft of the aperture  912 . Actuators  931  may be positioned along the corresponding flow surface  902  within the inlet duct  913  to energize the boundary layer flow, and in particular, to prevent flow separation as the flow turns into the inlet duct  913  from the forebody  915  and from regions aft of the inlet aperture  912  (indicated by streamlines  922 ). The orientation of the actuators  931  at the flow surface  902  can be spanwise and/or streamwise. In the streamwise orientation, the energy delivered by the actuators  931  can create a vortex that rolls up along the streamwise direction. The vortex can accordingly mix higher momentum air outside the boundary layer into the lower momentum boundary layer flow, delaying boundary layer separation. Optionally, the actuators  931  may also be installed at the forebody  915 . 
         [0046]      FIGS. 10A-10C  illustrate representative flow control assemblies installed on an airfoil to energize and/or control the adjacent boundary layer. Beginning with  FIG. 10A , a flow control assembly  1030  is installed on an airfoil  1020  having a flow surface  1002  with a leading edge  1021 . The flow control assembly  1030  can include actuators  1031  arranged in a single row at or near the leading edge  1021 . In another arrangement, the actuators  1031  can be arranged in a single row aft of the leading edge  1021 , as is also shown in  FIG. 10A . It is expected that in at least some embodiments, the single row of actuators  1031  (whether located at or near the leading edge  1021 , or well aft of the leading edge  1021 ) will be sufficient to energize the boundary layer passing over the flow surface  1002  in a manner that reduces skin friction and/or reduces the tendency for the flow to separate from the flow surface  1002 . 
         [0047]      FIG. 10B  is a partially schematic, cross-sectional illustration of the airfoil  1020 , illustrating one actuator  1031  from the row of actuators  1031  positioned at the leading edge  1021 . In this arrangement, the actuator  1031  is the only actuator  1031  carried by the airfoil  1020  at a streamwise line  1022  that intersects the actuator  1031 . By positioning the actuator  1031  at or near the leading edge  1021  (e.g., at or near the stagnation point), it is expected that the flow control assembly  1030  will improve the performance of the airfoil  1020  at angles of attack. In particular, it is expected that the flow control assembly  1030  will reduce the likelihood for flow separation near the leading edge  1021  at high angles of attack. 
         [0048]    In other embodiments, additional rows of actuators may be added to the airfoil, depending upon airfoil geometry and/or expected flight conditions. For example, as shown in  FIG. 10C , the actuators  1031  can be arranged in two or three rows proximate to the leading edge  1021 . The multiple actuators, spaced apart from each other in a streamwise direction, can create multiple mixing vortices that re-energize the boundary layer. An advantage of this arrangement is that vortical mixing is introduced in the streamwise direction and can persist to provide re-attachment and/or delay separation, not only in the immediate region around the actuators  1031 , but also downstream (e.g., well downstream) of this region. 
         [0049]    In any of the installations described above with reference to  FIGS. 3-10 , the actuator can have a dielectric barrier discharge configuration, generally as shown in  FIG. 1 , or the actuator can have other configurations that also add energy to the flow in a directed manner to produce a more favorable velocity distribution. The more favorable velocity distribution can produce enhanced performance, e.g., an increased region of laminar flow, a delay in boundary layer separation, an increase in flow uniformity, and/or an expanded flight envelope before separation is encountered.  FIG. 11A-12  illustrate embodiments of devices that add energy to the flow in accordance with mechanisms other than a dielectric barrier discharge mechanism, with the expected results being at least similar to those described above. 
         [0050]      FIG. 11A  illustrates a flow control assembly  1130  configured in accordance with one such embodiment. The flow control assembly  1130  can include an actuator  1131  that in turn includes at least two spaced-apart plates  1140  (shown as a first plate  1140   a  and a second plate  1140   b ). The two plates  1140   a ,  1140   b  are separated by a gap G and are coupled to a controller  1137 . In a particular embodiment, the controller  1137  applies a high voltage to the first plate  1140   a , with the second plate  1140   b  operating as a ground plane. The first plate  1140   a  can also include holes  1141  that allow particles (e.g., electrons) to exit the actuator  1131  into the adjacent flow. Accordingly, the actuator  1131  can operate in the manner of a hollow cathode array or particle accelerator. In particular, the voltage difference between the first plate  1140   a  and the second plate  1140   b  can accelerate electrons toward the first plate  1140   a . The electrons then exit through the first plate  1140   a  via the holes  1141  (as indicated by arrows E) and collide with air molecules in the adjacent flow. It is expected that in at least some modes of operation, the transfer of kinetic energy from the electrons to the air molecules can energize the adjacent boundary layer and can accordingly reduce the likelihood for the boundary layer to separate. 
         [0051]    In particular embodiments, it is desirable to obtain a high electric field between the two plates  1140   a ,  1140   b  so as to accelerate the electrons to a sufficient velocity to energize the adjacent boundary layer flow. For example, the first plate  1140   a  can have applied to it a voltage in the range of hundreds to thousands of volts, and the field strength can be increased by minimizing the gap G between the first and second plates  1140   a ,  1140   b . In a particular embodiment, the gap G can have a value in the range of from about 0.040 inches or less, the holes can have a diameter of about 0.040 inches or less, and the holes can be spaced apart to create a wide range of energy densities which are expected to be sufficient to generate a high energy electron stream, without causing the two plates to arc. For example, the holes can have a diameter of 75 micron or less. Suitable cathode discharge devices are described in U.S. Pat. No. 6,518,692, U.S. Pat. No. 6,528,947, and an article entitled “Development and Characterization of Micromachined Hollow Cathode Plasma Display Devices,” (Chen et al., Journal of Microelectromechanical Systems, October 2002) incorporated herein by reference. 
         [0052]    In particular embodiments, the flow control assembly  1130  can include multiple actuators  1131 . For example, as shown in  FIG. 11B , the flow control assembly  1130  can include an array of actuators  1131 , each having characteristics generally similar to those described above with reference to the actuator  1131  shown in  FIG. 11A . The actuators  1131  can be distributed over the surface of an airfoil  1120 , or any of the other surfaces and/or installations described above with reference to  FIGS. 3-10C . The actuators  1131  can be positioned over an entire forward region of the airfoil  1120 , or the actuators  1131  can be positioned only at selected locations over the surface of the airfoil  1120 . In any of these embodiments, the controller  1137  can individually control each of the actuators  1131 , and/or groups of the actuators  113 - 1 . For example, the controller  1137  can control the waveform parameters applied to each individual actuator  1131  and/or groups of actuators  1131 , e.g., the frequency, pulse width, amplitude, duty cycle, and/or phase relationship of the signals emitted by each of the actuators  1131 . In particular embodiments, the individual control over each of the actuators  1131  can allow the actuators  1131  to be operated in a synergistic manner. For example, the actuators  1131  can be controlled to create waves that constructively and/or destructively interfere with each other in a manner that can better manipulate the adjacent boundary layer flow. In a particular embodiment, the actuators  1131  can be spaced apart and activated in accordance with a schedule that produces constructively interfering (e.g., coalescing) shock waves. Suitable positions for the actuators  1131  and associated timing schemes can be determined experimentally using flow visualization techniques such as Schlieren techniques. In addition to or in lieu of such techniques, predictive tools (e.g., computational fluid dynamic or CFD tools) can be used to determine appropriate locations and timing patterns for activating the actuators  1131 . In any of these embodiments, synchronizing the activation schedules of the multiple actuators  1131  is expected to increase the strength of the resulting shock waves, which in turn is expected to increase boundary layer mixing and therefore reduce the likelihood for boundary layer separation. In particular embodiments, individual actuators  1131  can produce hemispherical shock patterns, and lines of actuators  1131  can produce hemicylindrical shock patters. These patterns can be selectively timed to constructively interfere with each other to establish enhanced boundary layer mixing in any of a variety of directions (e.g., a streamwise direction or a cross-stream direction). This arrangement can be particularly important because turbulence in such boundary layers is generally a three-dimensional effect (particularly when the airfoil  1120  is swept or forms a portion of a rotating rotor). By controlling individually addressable actuators  1131 , the controller  1137  can activate particular actuators  1131  at particular times, with particular energies and/or phase delays to produce a three-dimensional effect that is expected to more effectively control the adjacent boundary layer. The activation schedule for individual actuators  1131  and/or groups of actuators  1131  can be changed to account for changes in local flow conditions (e.g., freestream Mach number, angle of attack and/or others). 
         [0053]    In particular embodiments, the airfoil  1120  can include a detector  1139  (shown schematically in  FIG. 11B ) that in turn includes one or more detector elements positioned to identify flow characteristics. The detector  1139  can include one or more pressure taps, one or more hot film anemometers, and/or other devices that are located downstream from the actuators  1131  (as shown in  FIG. 11B ), within the array of actuators  1131 , and/or upstream of the actuators  1131 . The detector  1139  can be used to identify target flow conditions, e.g. incipient flow separation, or actual flow separation, turbulent flow, or laminar flow. These flow characteristics can change over different flight conditions, which can dictate a change in the number and/or location of actuators that are activated, and/or the wave form or other parameters in accordance with which energy is emitted by the actuators. Accordingly, based on information received from the detector  1139 , the controller  1137  can identify target actuators  1131  to activate, and can be programmed to deliver signals to the target actuators  1131  that are expected to enhance the characteristics or the adjacent boundary layers flow in a closed-loop feedback manner. 
         [0054]    The arrangement of actuators (e.g., hollow cathode arrays) shown in  FIG. 11B  can be applied to actuators other than the particle accelerator actuators shown in  FIG. 11A . For example, such an arrangement can be used in conjunction with the dielectric barrier discharge actuators described above with reference to  FIG. 1 . 
         [0055]      FIG. 12  illustrates an airfoil  1220  having a flow control assembly  1230  in accordance with still another embodiment of the disclosure. In this particular embodiment, the airfoil  1220  include opposing fences  1223  projecting generally normally outwardly from the aerodynamic surface. For purposes of illustration, the near-side fence  1223  (shown toward the bottom of  FIG. 12 ) is shown partially cut away. The fences  1223  can house opposing radiation emitters  1231  (e.g., laser emitters) that direct radiation across the surface of the airfoil  1220  to corresponding receivers  1242  (e.g., reflectors). Each of the emitters  1231  can emit radiation selected to correspond (or approximately correspond) to the absorption spectrum for air (e.g., laser radiation of wavelengths greater than about 200 nm). Each of the radiation beams emitted by the corresponding emitters  1231  can be oriented so as to have the desired direction across the surface of the airfoil  1220 . Adjacent emitters  1231  can be fired in succession to produce a desired effect (e.g., a traveling wave) on the adjacent boundary layer flow. The emitters  1231  can be positioned on opposite sides of the airfoil  1220  to increase the uniformity with which energy is directed into the adjacent airflow across the span of the airfoil  1220 . Each of the receivers  1242  can receive unabsorbed radiation from a corresponding oppositely-positioned emitter  1231  and reflect the received energy, dissipate the received energy, or recycle the energy so that it can be redirected into the adjacent boundary layer flow. Accordingly, the foregoing arrangement of emitters and reflectors can direct energy along multiple passes through the intervening space to increase absorptive heating. Suitable laser generators are described in an article titled “Process Study of a 200 nm Laser Pattern Generator,” (Hye-Keun Oh, Journal of the Korean Physical Society, December 2002), incorporated herein by reference, and are available from CryLaS GmbH of Berlin, Germany. 
         [0056]    In another embodiment, a flow control assembly can include a surface mounted array of UV light sources. The array of UV light sources can, in one embodiment, be arranged similarly to the flow control assembly  1130  of  FIGS. 11A and 11B . 
         [0057]    In other embodiments, the flow control assembly  1230  can have other arrangements. For example, rather than an array of fixed laser emitters and receivers, one or more of the laser emitters can rotate or otherwise be movable so as to sweep the emitted laser beam over the surface of the airfoil  1220 . In any of these embodiments, it is expected that the energy provided by the laser to the adjacent boundary layer flow will act to delay the transition of the flow from laminar flow to turbulent flow, and/or delay the location at which the boundary layer separates. For example, the energy emitted by the laser emitters  1231  can be pulsed in a manner generally similar to or analogous to that described above with reference to  FIGS. 2A-2B . 
         [0058]    From the foregoing, it will be appreciated that specific embodiments have been described herein for purposes of illustration, but that the disclosure encompasses additional embodiments as well. For example, the actuators described above may have configurations other than those specifically shown in the Figures. The signal parameters in accordance with which the actuators are activated may also have values other than those specifically shown and described above, and such values may be selected in a manner that depends upon the particular installation and/or the flight conditions to which the aerodynamic surface is exposed. The actuators may be installed on geometries, having features other than those specifically shown in the Figures. For example, the actuators may be applied to short, highly offset inlet diffusers. Certain aspects of the disclosure described in the context of particular embodiments may be combined or eliminated in other embodiments. For example, the actuators shown in  FIG. 1  may be combined to form an array generally similar to that shown in  FIG. 11B . Further, while advantages associated with certain embodiments have been described in the context of those embodiments, other embodiments may also exhibit such advantages, and not all embodiments need necessarily exhibit such advantages to fall within the scope of this disclosure. The following examples provide additional representative embodiments.