40th AIAA Joint Propulsion Conference AIAA-2004-4129 11 July – 14 July 2004 Fort Lauderdale, FL
Plasma Aerodynamic Flow Control for Hypersonic Inlets
David M. Van Wie* and Ashish Nedungadi† The Johns Hopkins University Applied Physics Laboratory, Laurel, MD, 20723
Plasma aerodynamics represents an emerging field involving the generation and utilization of plasmas for aerodynamic flow control. Current areas of active interest include volume discharges (including plasma jets, DC, HF, microwave and electron beam discharges) for large-scale inlet flow control, surface discharges (including DC, sliding and surface microwave discharges) for near-surface plasma flow control, large-scale and local MHD flow control techniques, and MHD power extraction techniques in hypersonic inlets. In addition, plasma aerodynamic and MHD flow control techniques are being investigated for controlling the flowfield in leading edge regions for the purpose of reducing drag and heat transfer. This paper provides a survey of plasma aerodynamic and MHD flow control techniques as applied to hypersonic inlets together with recommendations for promising areas of future research.
I. Introduction Plasma aerodynamic and magnetohydrodynamic (MHD) flow control technologies have been investigated intensely over the past decade for application to supersonic and hypersonic flight vehicles. Advances in the field are routinely discussed at the AIAA Weakly Ionized Gas Workshops, which are conducted as part of the Aerospace Sciences Meetings; AIAA Plasmadynamics and Lasers Conferences; the series of Workshops on Magnetoplasma Aerodynamics for Aerospace Applications sponsored by the High Temperature Institute of the Russian Academy of Science; and the series of Workshops on Thermochemical and Plasma Processes in Aerodynamics sponsored by the Leninetz Holding Company in St. Petersburg, Russia. In these techniques, plasmas are created within flowfields for flow control or energy extraction purposes through physical processes, such as energy deposition or generation of electrostatic or electromagnetic forces on the flow. A number of survey papers have been generated summarizing the individual technologies1-5 and a preliminary system assessment of many of these technologies was presented in Ref. 6.
Many aspects of the technologies under consideration have been investigated for several decades, but the recent resurgence in their interest can be traced to the AJAX vehicle concept, which was developed by the Leninetz Company to support either long-range hypersonic cruise or space access missions7,8. The AJAX vehicle concept incorporates plasma aerodynamic flow control, MHD flow control, power generation, and flow acceleration, and an endothermic fuel reforming process in combination with a basic hydrocarbon-fuel scramjet engine. The AJAX-vehicle technologies provide intriguing system-level synergies for enhancing overall vehicle performance, but the individual technologies also offer the potential for solving a wide range of design problems for hypersonic vehicles. In this study, the individual technologies are addressed with respect to their potential applications to solving design challenges associated with scramjet inlet design and operation are considered.
Prior to discussing the application of plasma technologies, fundamental scramjet inlet performance and operability issues are discussed. Recognizing that the inlet of a scramjet engine serves to capture and compress the engine airflow prior to introduction of fuel within the combustor, one can state the primary goal in the development of any scramjet inlet system is to define a geometry that provides an efficient compression process, generates low drag, produces nearly uniform flow entering the combustor, and provides these characteristics over a wide range of flight and engine operating conditions. A number of the principal inlet performance and operability challenges are presented schematically in Fig. 1. Performance issues exist associated with blunt leading edge effects, boundary
layer transition, skin friction and heat loss to the walls, supercritical spill at speeds below the design Mach number, and shock losses through the compression process. Operability issues exist associated shock-shock interactions at the cowl leading edge, shock-wave/boundary-layer interactions at shock wave impingement locations, and isolation of the pre-combustion pressure rise within the inlet isolator.
A variety of “conventional” approaches have been investigated for solving many of the design issues associated with the inlet issues identified in Fig. 1, but there are system complexities and compromises inherent in these approaches. For example, inlets can be divided into fixed- and variable-geometry classes. Fixed-geometry designs offer a lightweight approach by using a point-design while attempting to provide acceptable performance over a limited range of Mach number and angle-of-attack. In vehicle designs where the performance of a fixed geometry design is not acceptable, the weight, cost and complexity of a variable geometry design is required to provide adequate performance over the required Mach number and angle-of-attack operating range. The plasma-based techniques discussed below potentially offer better approaches to address these inlet design challenges. A compilation of the plasma-based techniques with application to inlet design is shown schematically in Fig. 2. These techniques can be categorized in terms of MHD-based techniques for energy management engines, large-scale inlet flow control techniques, near-surface flow control techniques, and leading edge flow control techniques. MHD-based techniques for energy management have been suggested as a means for producing large levels of on-board power and for modifying the basic thermodynamic cycle of the engine. Large-scale inlet flow control techniques have been suggested for shock position control in the inlet, shortening the isolator requirement, or acting as a virtual cowl. Each of these techniques is aimed at improving the performance of the inlet or reducing the mass of the engine by shortening the engine or minimizing the variable geometry requirements. Near-surface flow control techniques have been suggested for reducing skin friction losses, controlling boundary layer transition and shock-wave/boundary-layer interactions, and controlling internal duct shock position. Leading edge flow control techniques have been suggested for minimizing drag and heat transfer and disrupting the shock-shock interactions than can occur at the cowl leading edge. In the following sections, sample flow control techniques will be discussed in each category.
II. MHD-Based Energy Management Engines An MHD power generation system offers the potential to extract significant levels of power from high-speed electrically conducting streams. The possibility of using an MHD energy extraction system to fundamentally alter the thermodynamic cycle of the basic scramjet engine is an inherent feature of the original AJAX system concept7-11. This concept has been explored by a number of research teams and expanded to include alternate MHD-based energy management schemes. As illustrated in Fig. 3, these schemes can be divided in categories depending on where the energy extraction occurs and for what purpose the extract power is used. The original AJAX concept consists of energy extraction in the inlet with the energy bypassed around the combustion process and re-introduced in an MHD-acceleration system. An alternate approach is to extract energy downstream of the combustor and use this energy in the inlet using one of the flow control techniques. A third and fourth approach is to generate the energy in an auxiliary MHD power system, either using on-board propellants to fuel a rocket thruster with an MHD duct or using energy extracted directly from the external airstream, and then use the energy in one of the flow control techniques. Rather than discussing the system implications of these energy management approaches, this paper will focus on the inlet implications of energy extraction and flow control.
MHD energy extraction within an inlet requires acceptable levels of electrical conductivity to support the magnetic interaction. As illustrated in Fig. 4 for an optimum 4-shock scramjet inlet, the static temperature in the inviscid flow of inlets remains below approximately 2000K for flight speeds as high as Mach 20 at flight dynamic pressures between 0.24 and 0.96 atm. With these low static temperatures, the naturally existing electrical conductivity of the air will be insufficient to sustain a significant magnetic interaction, so an active flow ionization technique must be used. The most energy efficient technique for providing sufficient conductivity is through the use of electron beams although a repetitively pulsed high-voltage discharge may prove acceptable12-15.
Macheret et. al.16 investigated the operating characteristics of an MHD inlet power extraction system that would be located in the forward portion of a vehicle on flow turning surfaces at low angles with an emphasis on extraction of the maximum possible energy for bypass applications. This work focused on the use of an electron beam system for generating adequate conductivity with high magnetic fields and relatively low power requirements for the ionization system, which resulted from operating at a forward location in an inlet where low flow turning angles exist. This class of energy production system resulted in a long power generation system and high Hall parameters. Calculations were conducted for an MHD channel with an inflow cross-section of 25 x 25 cm2 and a length of 3 m. A magnetic field of 7 Tesla was assumed. Results indicate that 8-12% of the flow kinetic energy could be converted
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to electrical power at low altitudes at Mach numbers between 4 and 10, while up to 33% could be converted at high altitude. Kuranov et. al.9-11 investigated the use of nonequilibrium ionizations sources for an inlet-based energy extraction system as part of the AJAX concept and demonstrated power extraction levels consistent with that required for altering the thermodynamics of the engine cycle. Slavin et. al.17 investigated the use of nonuniform gas-plasma flows for energy extraction onboard a hypersonic engine where they showed energy extraction levels up to 30% at a Mach 7, 30-km altitude condition. These techniques generally show significant levels of power generating capacity although significant losses of total pressure can be realized when the power is extracted as part of the inlet. Significant variations exist in the assessment of the MHD power extraction and its subsequent on the overall propulsion system performance9-11, 18,19.
A potentially advantageous byproduct of an MHD energy extraction system in the inlet concerns the flow profiles exiting the generator and the chemical state of the airstream. MHD power generators to produce a more uniform flow profile at their exit, so the generator may positively condition the flow from the viewpoint of uniformly fueling the engine in the combustor. In addition, the exhaust of the MHD generator will likely not be in chemical and/or thermal equilibrium, which will impact chemical reaction rates within the combustor. Given the nonequilibrium condition of the gas exiting the MHD power generator, enhanced ignition and flameholding characteristics may be realized.
III. Large-Scale Inlet Flow Control Large-scale inlet flow control techniques include MHD-based systems for shock position control and plasma
volume discharge techniques aimed at modifying large-scale features of an inlet flowfields. The four main techniques for large-scale inlet flow control are illustrated in Fig. 4. The first two techniques have been suggested as techniques where scramjet inlets can potentially benefit from MHD flow control, especially in situations where the inlet is required to operate over a significant range of Mach number and/or angle-of-attack. MHD flow control for scramjet inlets has been suggested for use in control of the captured flow rate, control of shock positioning, control of compression ratio, and for control of local shock-wave/boundary-layer interactions20-28. The first concept (Fig. 4a) uses the MHD inlet control system to decelerate the inlet flow and reposition over-sped shocks back onto their design condition, such as shock-on-lip operation. This concept could potentially be used on a vehicle designed for operation at low hypersonic speeds; hence, the inlet is designed for low speed and then the MHD system is used to maintain shock position control as the vehicle is accelerated to higher speeds. The second concept (Fig. 4b) is used to add additional compression to the inlet flow using MHD flow control to draw additional mass through the engine at under-speed conditions. Hence, the inlet is designed for high-speed operation and then the MHD system is used to control the flow at lower speeds. Little information is available on this promising mode of operation. The third concept uses a plasma discharge to create a virtual cowl29, which deflects additional mass into the engine at under-sped conditions. The final concept (Fig. 4d) uses plasma discharges to alter the characteristics of the isolator30.
Large-scale inlet flow control techniques have been investigated experimentally and analytically. Experiments aimed at demonstrating MHD flow control in a scramjet inlet were conducted in the “small shock tunnel” of the Ioffe Physico-Technical Institute31-35. In these experiments the issues associated with creating and maintaining an ionized airstream were separated from the basic problem of flow control, by testing in a shock tunnel using noble gas plasmas. As illustrated in Fig. 6, the shock tunnel consists of a 1 m long high-pressure driver tube, a 4.325 m-long low-pressure driven tube coupled to a 0.125 m long measurement tube, two-dimensional converging-diverging supply nozzle, test chamber, and vacuum tank. The diameter of the shock tube is 5 cm with a facility run time of approximately 400 µs. To achieve sufficiently high electron number densities in the freestream, rare gases were used as the test media since the recombination and electron attachment rates are much lower compared to air. For the majority of the testing, the shock tube was operated with krypton as the test gas. After shock heating the driven gas to high temperatures where equilibrium thermal ionization is achieved in the reservoir region, the gas is passed through the facility converging-diverging nozzle where the expansion is sufficiently rapid that the electron recombination does not remain in equilibrium with the expanding gas. Operating with a driven Mach number of 7.8 and an initial driven tube pressure of 20 Torr, the temperature, density, and conductivity of the Krypton in the reservoir were estimated to be 9800K, 1.08 kg/m3, and 3200 S/m. Expanded the gas through the nozzle to Mach 4.2 conditions, the temperature, density, and velocity were estimated to be 1550K, 0.0645 kg/m3, and 2100 m/s. The resulting freestream gas has a sufficiently high electron number density to investigate the MHD interactions within the inlet flowfield. The inlet model consisted of a two-dimensional opposed-ramp design positioned at the downstream end of the facility supply nozzle as shown in Figure 7a. The supersonic portion of the nozzle and the inlet model are constructed of an insulating acrylic material with flush-mounted segmented copper electrodes spanning the width of
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the nozzle and inlet. Note that two pairs of electrodes were situated in the aft end of the facility nozzle, which were used to minimize the end effects when operated as a Faraday channel. The cross-section of the inlet at its entrance is 32 x 38 mm2. Each wall converged with a 5.5° angle over a length of 65 mm such that the throat cross-sectional area is 19.48 x 38 mm2, which translates to an internal contraction ratio (Ath/Ain) of 1.64. A photograph of the inlet model mounted in the test section is seen in Figure 7b. The surface mounted electrodes in the inlet are visible between the electric return lines, which run along the outside of the test section. The natural luminosity of the flowfields under test conditions are shown in Fig. 8 Since the luminosity is closely tied to the local density, the shock patterns through the inlet are apparent. The very luminous region along the lower surface of the photograph resulted from an overexposure of the film, but the remaining shock patterns can be discerned. As the magnetic field strength is increased, the shock waves steepen and the point of intersection moves forward in the inlet. At a field strength of 1.2 T, the shock waves in the inlet are seen to be nearly normal on the centerline of the inlet indicating a very strong interaction. Schlieren photographs of the flowfield are provided in Fig. 9 for magnetic field strengths of 0 T, 0.85 T, and 1.2 T. The results show that at a field strength of 0.85 T, the crossing shocks area steepened and moved forward in the inlet duct compared to the condition without the magnetic field such that the impingement point of the shock wave on the walls moves from downstream of the shoulder to a point upstream of the shoulder. At a field strength of 1.2 T, the character of the inlet shocks change significantly with multiple strong waves present in the flow. Measurement of the axial position of the crossing shock waves provides an estimate of the flow control exhibited in the inlet. Xc is defined as the axial position of the oblique crossing shocks when the shocks remain straight. For stronger interactions with near-normal shocks present, Xsh, is used to define the forward-most point of shock interaction. Measurements of Xc and Xsh are provided in Figure 10 for tests conducted at field strengths between 0 and 1.4 T. The squares designate measurements made from the high-speed movies of the natural luminosity and the circles indicate measurements obtained from the Schlieren pictures. The results show roughly three zones of behavior. For magnetic field strength below 0.7 T, the crossing oblique shocks remain well behaved as they move forward in the inlet in the interaction zone. This type of interaction has been designated a weak interaction. For field strengths between 0.7 T and 1.1 T, the point of shock crossing was unsteady in the high-speed photographs indicating an unstable region of operation. For field strengths above 1.1 T, the flowfield is again steady as indicated by the high-speed photographs, but the interaction contains near-normal waves in the inlet. This type of interaction is designated a strong interaction. The generation of the unstable mode of operation is believed to be tied to flow choking in the throat of the inlet. Establishment of the flowfield through an inlet with choked flow is known to take much longer to establish than that required to establish supersonic flow throughout the inlet, so adequate test time may not exist in this facility to fully explore this mode. Given sufficient time, the unstable shock system will likely travel forward of the leading edge of the inlet resulting in large flow spillage around the inlet. In the strong interaction mode, the flow is driven to subsonic speeds through the interaction zone in a manner that stable flow through the inlet results. The calculated Mach number distribution through the inlet without electric or magnetic fields is shown in Figure 11. This calculation was started near the throat of the facility supply nozzle, so that the effects of a non-uniform freestream could be captured. The results indicate that the flow is compressed from the Mach 4.2 freestream conditions to approximately Mach 2.8 at the inlet throat. The inlet is constructed such that the crossing shock waves intersect the model walls downstream of the inlet throat. Sample computational results are shown in Figure 12 for a case with discrete electrodes and an interaction parameter of St=0.6 with an applied voltage of 18 Volts across the channel. Contour plots for Mach number, static pressure, and potential are provided together with a vector representation of the current distribution. The computational results for this case of a weak interaction show that the crossing oblique shocks are drawn forward and cancelled on the inlet shoulder. The flowfield is well behaved although a complicated current distribution is observed even for this relatively simple scheme. Thus, the Ioffe Institute results show that with sufficient levels of magnetic interaction, inlet shock position control can be used to maintain an inlet operating at off-design conditions in rare gas plasmas. Significant levels of flow control were demonstrated including the ability to drive the flow into both unstable and strong interaction modes of operation. Analytic investigations into the creation and maintenance of air plasmas by been conducted using electron beam sources11,36 and results have demonstrated that sufficient levels of flow conductivity can be generated for large-scale flow control at power levels below that which can be extracted from the MHD system (i.e. a self-sustaining mode of operation). The two principal problems associated with the large-scale MHD flow control system is the total pressure losses associated with the MHD flow control system and the mass of the large magnet required. In Ref. 6 results were presented from a system assessment of a large-scale inlet flow control on a hypersonic vehicle. After conducting investigations into the preferred location of the MHD flow control system, results showed that the
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system should be located as far forward as possible and use a narrow intense interaction region. As illustrated in Figure 13, the concept called for a large torriodal magnet capable of generate 3 T at the vehicle surface to be integrated into the forward end of an inlet with electron beam generators positioned in the core of the magnet to provide the non-equilibrium ionization source. Results of two-dimensional CFD calculations with the MHD interaction are shown in Figure 14 for the case with a Mach 5 inlet design operating at Mach 8. Contour plots of Mach number, static pressure, static temperature, and electron number density are provided. The narrow region of MHD interaction is clearly visible in the electron density contour plot, and the repositioning of the forebody shocks at the location of a cowl lip with near shock-on-shoulder operation can be seen in the other contour plots. Thus, the calculations indicate the large-scale inlet flow control can be accomplished in air using non-equilibrium ionization.
The impact of the MHD flow control system on the overall engine operation was examined in Ref. 6 using a generic scramjet engine as a baseline Mach 10 design. The inlet performance with the MHD system was calculated using 2D CFD and fed into a quasi-1D engine cycle analysis. The ratio of the engine specific impulse, Isp, with the MHD flow control system to the baseline engine Isp is shown in Fig. 15 for Mach numbers between 5 and 10. The results show that the large-scale MHD flow control system results in significant engine Isp penalties as the speed is increased. The second problem associated with the large-scale flow control system in the large mass of the magnets. In Ref. 6, superconducting electromagnets were assumed, yet the magnet mass overwhelmed any potential savings in fuel. The potential for significant reductions in magnet mass37 and improvements in cycle integration exist and need to be pursued before large-scale MHD flow control systems will become feasible.
Alternate technique for creating large-scale inlet flow control techniques have been aimed at plasmadynamic discharges in the flowfield to create disturbances that modify the inlet flowfield in an advantageous manner. The concept of the virtual cowl29, relies on energy deposition upstream and slightly outside of the cowl lip to deflect mass flow into the engine, thus provide some of the features of a mechanically variable inlet. In Ref. 29, calculations were conducted to demonstrate that the mass flow of an inlet operating at speeds below its design condition could be significantly increased (up to 33% at Mach 6 with 10% enthalpy extraction for a Mach 10 inlet design) using the virtual cowl concept. Using an MHD power extraction system downstream of the combustor to provide the power needed to create the plasma discharge, the effect of the virtual cowl system is principally one of increasing the thrust potential of the engine by forcing through more airflow. In Ref. 6, calculations of the impact of the virtual cowl on a generic baseline engine operation showed a negligible effect on specific impulse with the engine thrust nearly proportional to engine mass flow. The one significant disadvantage of the virtual cowl concept is that significant levels of power are required (MWs/m of engine width), so the mass associated with the power generation system is large.
The final large scale inlet flow control technique to be discussed is associated with the use of plasma discharges to shorten the isolator, which tends to be longer, heavier, and require significant cooling at high hypersonic speeds. Remembering that the isolator is used to contain the shock system created by the heat release in the combustor, the isolator must be sufficiently long to contain this pressure rise, and this required length increases significantly as the inflow Mach number increases. Shneider et. al.30 have investigated the use of plasma flow control in an isolator and has shown that the Mach number downstream of the plasma discharge can be reduced significantly, which will result in a shorter isolator requirement. Using distributed heating in the internal portion of the inlet with heating rates in the range of 6-8% of the total enthalpy, the flow Mach number could be decreased to nearly sonic conditions, which may avoid the need for an isolator. The losses associated with this heating and the power extraction, if done downstream of the scramjet combustor, combine to reduce the engine thrust by about 16% at a Mach 5 condition. Detailed system assessments will be required to assess the impact of a shorter isolator with its reduced mass and thermal cooling requirements on the overall system performance. While the preliminary performance estimates appear interesting, significant questions exist concerning the stability of the process and the coupling that would occur between the inlet and combustor. Additional efforts will be required to understand the control aspects of such a system.
IV. Near Surface Plasma Flow Control Techniques At hypersonic speeds, inlet efficiency losses due to skin friction represent a significant fraction of the total
inefficiency that accrues in the compression process. In addition, uncertainties in the location of boundary layer transition, and the material and cooling requirements associated with a stressing aerothermal environment place a premium on potential techniques capable of reducing skin friction and heat transfer. While the large-scale inlet flow-control techniques require high power levels to maintain flow conductivity or deposit sufficient energy for flow control, a number of near-surface plasma flow control techniques have been investigated which offer the potential to favorably impact inlet performance with much smaller power requirements. Surface discharge techniques that have
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been recently investigated for supersonic and hypersonic aerodynamic applications include surface microwave discharges38-40, DC and pulsed discharges between surface mounted electrodes41-43 and in vented volumes44, and sliding discharges45. These techniques offer the potential for rapidly and selectively heating a boundary at or slightly off the surface and/or creating zones of weakly ionized gas for MHD interactions. MHD flow control techniques have also been investigated for affecting boundary layer transition and modifying the basic characteristics of turbulent boundary layers46-51.
Leonov et. al.41,42 have investigated the use of a discharge between surface mounted electrodes to reduce the friction drag of a flat plate. At transonic speeds with a local static pressure of 500 Torr, a 15kW discharge was shown to result in gas heating sufficient to produce a gas temperature of approximately 1500K. A portion of the flat plate downstream of the discharge was isolated and mounted on a force balance. The drag measurements provided evidence that the friction drag downstream of the discharge is significantly reduced. Microwave surface discharges have also been investigated for the purpose of modifying the boundary layer characteristics in flowfields. Shibkov et. al.38-40 have demonstrated the capability to generate and stabilize a discharge in the presence of a Mach 2 airstream over a length of 15 cm with a cross-sectional area of 1x2 cm2. The time variation of the temperature in the discharge has been measured as a function of input power with initial heating rates on the order of 50K/µs. The very fast gas-heating rate in discharges with high reduced electric fields (E/n>100 Td) has been attributed to the effective excitation of electronically excited states of nitrogen and their subsequent fast quenching. As the pulse power is increased from 35 kW to 175 kW, the gas temperature was found to increase from 500K to 1700K while the vibrational temperature remained nearly constant. Thus, microwave surface discharges provide a means for rapidly heating a boundary layer while a fine control on the resulting temperature can be maintained.
Surface discharges have been shown to be an effective technique for adding thermal energy to a boundary layer. For speeds below approximately Mach 6, the creation of a hot layer near the wall that stabilizes the boundary layer so that transition can be delayed. Above Mach 6, alternate transition modes become dominant and the effects of a discharge are not clear. Levin et. al.52 have calculated the reduction in turbulent surface skin friction coefficient on a flat plate at Mach 3 due to a surface discharge, which was modeled as an energy deposition in a small volume above the surface. The discharges were shown to impact the skin friction well downstream of the discharge region and energetic efficiencies approaching 50% were calculated for the Mach 3 conditions. No information is available on the effect of increasing Mach number on drag reduction energetic efficiency of surface discharges, and the impact of surface discharges and their resultant decrease in integrated skin friction in inlet efficiency have not been addressed.
In addition to the use of plasma discharges for heating the boundary layer near the surface, surface discharges that create a weakly ionized gas can be used in an MHD flow control concept. Two potential schemes for local MHD flow control in an inlet have been proposed. The 1st concept is an inlet laminar flow control system that extends from the observation that boundary layer transition in most inlets can be traced to either Görtler instabilities created in the concave turning of the compression surface or instabilities in the separated shear layer resulting from laminar boundary layer separation. As illustrated in Fig. 16, a localized MHD-flow control system can potentially be used to stabilize these transition sources and extend the laminar region in the hypersonic inlet. Sample engine performance calculations were conducted to assess the impact of a laminar flow inlet for Mach numbers between 6 and 12 for flight at dynamic pressures of 0.47 and 0.95 atm. The resulting engine Isp and CT are shown in Fig. 17 where modest improvements are seen. Potentially more significant results from the inlet laminar flow control system concern the mass of the thermal protection system, which would be lower with the laminar flow control system.
The second concept uses local MHD flow control for shock-wave/boundary-layer interaction (SWBLI) control as illustrated in Fig. 18. Conventional inlets are often designed for operation against SWBLI constraints to prevent boundary layer separation and the resulting shock-expansion losses and unsteady flow features. Attempts at incorporating bleed and blowing systems to stabilize hypersonic SWBLI phenomena have generally not been successful. With local MHD flow control, it is theorized that boundary layer separation can be avoided by “pumping” the low-momentum flow near the wall through the adverse pressure gradient of the shock interaction. Compared to the global MHD flow control techniques, this local technique has the advantages that the local velocity is low, the boundary layer temperature is high, and the magnetic field and ionized regions are small.
The advantage of a local MHD SWBLI control system was evaluated for an engine concept that employed a rotating cowl. In the baseline engine, the performance of the engine was limited by a “weak-shock” (WS) constraint, which is a boundary layer separation criterion. The results of engine performance calculations with and without the MHD flow control system are shown in Fig. 19. In this technique, the Lorentz force is oriented to accelerate the low-momentum flow near the wall through the adverse pressure gradient of the shock interaction. With the MHD flow control system, the engine thrust can be significantly increased at Mach numbers below the design operating condition since the engine cowl can be rotated outward and additional airflow captured. For speeds
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at or above the design Mach number, the effect of the cowl rotation is minimal as the cowl lip captures low-density uncompressed freestream air. Also note that the engine specific impulse is only slightly changed by the use of the local MHD flow control system. With the added thrust and small weight penalties of the local MHD flow control system, this technique shows promise for a positive impact on the overall hypersonic system performance.
Attempts at experimental demonstrations of MHD SWBLI control are underway to the Ioffe Physical Technical Institute53-57 and JHU/APL. At the Ioffe Institute, in their “Big Shock Tube”, the test equipment is designed for studying the influence of MHD effects on a supersonic flow of a weakly ionized gas using rare gas and molecular plasmas. At the nozzle inlet, xenon plasma heated by incident and reflected shock waves and has a concentration 8⋅1019 cm−3 and temperature of about 9000 K. The flow of duration is approximately 2 ms with the velocity corresponding to Mach number 5 at the outlet of a planar supersonic nozzle. The objects under study are the shock wave structure of the flow and electrophysical properties of the plasma flow at a 15° compression corner, which is formed by a flat plate attached to the nozzle wall at the nozzle outlet (Fig. 20). The flow structure is analyzed with the use of the Schlieren method to visualize the oblique shock wave formed as a result of the flow turn inside the compression corner, and to watch variations of its position and shape. In the walls of the nozzle and deflecting plate near the compression corner vertex two pairs of electrodes are embedded right up to the window glasses. Through each pair of electrodes an electric current can be passed transverse to the plasma flow. An external magnetic field with a duration of 4.5 ms and the maximum magnetic induction 1.5 T is directed normal to the plasma flow and the electric current direction. The synchronization of the equipment operation with the process under study is performed with the help of piezoelectric transducers located in the shock tube channel. Photometric and electric measurements provide detecting the radiation and conduction of the plasma.
The initial experiments have been conducted using a decelerative mode of the MHD interaction, which will be unlikely to impact positively the boundary layer separation limits. Schlieren photographs of the corner flow region are shown in Fig. 21 for cases with the current drawn through the upstream and downstream electrodes. In Fig. 21a, the upstream current flow has little impact on the shock wave emanating from the corner. When the downstream electrodes were energized, a splitting in the shock was observed (Fig. 21b). This splitting is due to a curvature introduced in the shock wave with the portion of the shock near the test section walls being much more influenced than at the centerline of the experiment. As the magnetic interaction is increased, this splitting between the near-sidewall and centerline shock wave increases as shown in Fig. 22. When the magnetic interaction parameter (or Stuart Number, St) is increased beyond 0.6, flow separation was created at the corner as shown in Fig. 23.
A new experimental facility is currently being assembled at JHU/APL to investigate the potential use of local near-surface MHD control for expanding shock-wave/boundary-layer interaction limits within an inlet. As illustrated in Fig. 24, supersonic flow will be provided by a Laval nozzle to a test section that passes through an electromagnet. A shock generator will be used to create an oblique shock that will impinge on the tunnel wall boundary layer in the interaction zone. Various discharge techniques will be investigated and an accelerative mode of MHD interaction will be explored. Testing is planned for the late summer, so preliminary results will be available later this year.
Leading Edge Flow Control Techniques Blunt leading edge effects in inlets at both the inlet leading edge and cowl lip become much more prominent at
hypersonic speeds. For large hypersonic vehicles, the fraction of the total vehicle drag due to the leading edges may be small, but the curved bow shock and resultant entropy layer can impact the entire inlet flowfield. The conventional approach to minimizing the impact of a blunt leading edge is to investigate novel materials and cooling strategies to allow relatively sharp leading edges to be used. Plasma discharges in front of blunt leading edges have been investigated for the purpose of modifying the flowfield around the leading edge and decreasing the leading edge drag58,59. Plasma discharge techniques that have been investigated include microwave discharges, high frequency discharges and plasma jet injection techniques. Plasma jet injection, wherein plasma is created internal to a model and injected into the flowfield for the purpose of modifying the flowfield and reducing the leading edge drag, has been investigated by a number of researchers59-63. Krasilnikov et. al.59-62 investigated the drag reduction potential of plasma jet injection using a cone-cylinder model. At speeds between Mach 0.4 and 5.0, the plasma jet injection was seen to reduce the drag by a factor of approximately two. Shang et. al.63 has investigated the plasma jet injection phenomenon in a series of experiments where the plasma jet injection was compared to conventional fluid injection. Experiments were conducted at Mach 6 using a plasma injector operating with a temperature of approximately 4400K, electron temperature of 20,000K and an electron number density of 3x1012/cm3. Plasma jet injection from a spherically blunted cylinder was shown to reduce drag although unsteady bifurcated flows can result over significant portion of the operating regime. The drag reduction aspects of the plasma jets appear to be
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principally fluid dynamic although a 12% difference in drag was observed between measured with a plasma jet and that expected with injection of a hot fluid at the same temperature. The plasma jets offer the advantage of operation with a high temperature injectant, which allows a given effect to be realized for a smaller amount of injectant. Interestingly, the plasma jet injectors also show lower pressure oscillations compared to corresponding cold fluid injectant cases in domains where the bifurcated flow results.
Microwave energy deposition upstream of blunt bodies has also been investigated for drag reduction applications64,65. Microwave techniques suffer from the requirement to incorporate a RF-transparent window into the inlet, but offer the advantage of being able to locate the discharge at some distance from the leading edge. Both pulsed64 and steady-state65 discharges have been investigated, and significant levels of drag decrease have been measured at speeds of Mach 1.5-2 with high energetic efficiencies.
While both the plasma jet injection techniques and the microwave discharge techniques have been demonstrated to lead to drag reduction with high energetic efficiencies, the implications of the use of this type of system on the operation of an inlet have not been investigated.
One of the challenging issues associated with the development of hypersonic inlets concerns the Edney Type IV shock-shock interaction that occurs when a leading edge shock intersects the cowl lip bow shock. In the Type IV interaction, a narrow jet is created at the intersection of the shocks, which impacts on the leading edge to create a zone of high heat transfer. Amplification factors of 40 over the undisturbed stagnation point heating rates have been measured. Plasmadynamic discharges have been investigated as a means of disrupting the Type IV interaction and significantly reducing the maximum heat transfer that a leading edge sees. Aldegren et. al.66,67 demonstrated a disruption of the Type IV interaction using a pulsed laser system to create a plasma discharge upstream of the shock-shock intersection location. The pulsed discharge was seen to result in a rapidly expanding wave system emanating from the region of the energy deposition, which impacts the shock-shock intersection and disrupts the establishment of the impinging jet. This technique appears promising and merits additional research including efforts associated with development of sensing and control systems.
In addition to plasmadynamic discharges, the use of MHD flow control for reducing heat transfer at leading edges has been investigated68-71. The concept of using MHD effects to reduce heat transfer dates to early hypersonic studies conducted in the 1950s, and a good review of this early effort is provided by Poggie and Gaitonde68. In this work, results are also provided from computational and theoretical studies of the hypersonic flow over a hemisphere, which were conducted to investigate the potential use of MHD effects for mitigating the heat transfer at a leading edge. Results show that the stagnation point heat transfer could be reduced by approximately 25% at Mach 5 with a magnetic interaction parameter of 6. Batenin et. al.69 provided calculations to indicate the potential reduction in heat transfer at a blunt leading edge using current flows along the leading edge to produce a circumferential magnetic field. The resulting Bj
rr× force leads to a deceleration of the flow approaching the leading edge. Analytic results
show that the peak heat transfer rate at the leading edge can be reduced in the presence of a magnetic field. Bityurin et. al.70,71 have conducted experiments on MHD flow interactions at leading edges in a Mach 12 to 15
conditions in an MHD driven wind tunnel. By passing electric current through the simulated leading edge, a circumferential magnetic field was generated with 1 T at the surface. While pre-test predictions of the interaction showed a significant impact on the flowfield due to the MHD interaction, preliminary experiments have shown only modest changes to the flowfield. The discrepancies between predictions and experiments were attributed to a dominating role of the Hall effect.
V. Conclusions Plasma aerodynamic and MHD flow control techniques are being aggressively investigated for application to
supersonic and hypersonic vehicles. These techniques offer advantages over “conventional” flow control techniques in that the control can often be more precisely placed and operated with response times much faster than existing electromechanical and hydraulically actuated systems. In addition, these flow control techniques may ultimately be useful for reducing variable-geometry requirements in an inlet, which are complex and heavy and result in difficult sealing issues at hypersonic speeds.
A wide variety of flow control techniques that have application to hypersonic inlet design have been investigated and the ability to control flow phenomena has been demonstrated. These techniques can be broadly categorized as large-scale flow control techniques, near-surface plasma flow control techniques, and leading edge flow control techniques. In addition, the potential exists for extracting power of an inlet at hypersonic speeds using an MHD power extraction system.
At the present time, the large-scale inlet flow control techniques have been demonstrated experimentally in rare-gas plasmas and computationally in air plasmas. Significant degrees of flow control are available although the
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power levels are large and the potential to significantly degrade overall engine performance exists. Additional work is required to decrease the mass associated with subsystem components and to favorably integrate the flow control into the overall engine cycle.
Near-surface plasma flow control techniques have been investigated for reducing skin friction, modifying the turbulence in boundary layers, and preventing shock-wave/boundary layer interactions. These systems have the advantage that the mass and volume associated with the flow control system is modest, so demonstration of a system performance improvement is easier. At the present time, preliminary computations and experiments have demonstrated interesting levels of flow control, but no overall advantageous impact on an inlet flowfield has been quantitatively demonstrated. Additional research work is needed to understand the physical processes, optimize the flow control systems, and assess the overall impact on an inlet and engine.
The use of plasma flow control techniques at leading edges has been investigated for reducing drag and heat transfer and disrupting complex shock-shock interactions. Experiments and computations have demonstrated significant levels of flow control and drag reduction. Preliminary experimental investigations of MHD flow control for heat transfer reduction at hypersonic speeds have shown modest effects. Additional research is needed on optimizing the flow control techniques and integrating the technique into realistic inlet configurations.
VI. Acknowledgments This activity was conducted under the Joint NASA/DoD University Research, Engineering, and Technology
Institute for Reusable Launch Vehicles. The NASA portion of the program was conducted under Next Generation Launch Technology program with Claudia Meyer as the Program Manager. The DoD portion of the program is overseen by AFOSR with Dr. John Schmisseur as program manager. At the Ioffe Physico-Technical Institute, the MHD flow control experimental program is headed by Prof. Sergey Bobashev and the computational program is headed by Prof. Yurii Golovachov. The large-scale inlet flow control scheme shown in Fig. 13 was generated by J. Silkey of The Boeing Company. The CFD Results shown in Fig. 14 were generated by Drs. S. Macheret and Shneider at Princeton University.
VII. References
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29. “Scramjet Inlet Control by Off-Body Energy Addition: A Virtual Cowl,” S. O. Macheret, M. N. Shneider, and R. B. Miles, AIAA-2003-0032, Jan. 2003.
30. “Modeling of Plasma Virtual Shape Control on Ram/Scramjet Inlet and Isolator,” M. N. Shneider and S. O. Macheret, AIAA-2004-2662, June 2004.
31. “Influence of MHD Interaction on Shock Wave Structure in Supersonic Diffuser,” S. V. Bobashev, E. A. D’yakonova, A. B. Erofeev, T. A. Lapushkina, V. G. Maslennikov, S. A. Poniaev, V. A. Sakharov, R. V. Vasil-eva, and D. M. Van Wie, 2nd Workshop on Magneto- Plasma-Aerodynamics in Aerospace Applications, Moscow, April 2000.
32. “Numerical Investigation of MGD Flows in the Models of Supersonic Intakes,” Yu. P.Golovachov, A. A. Schmidt, and S. Yu. Suschikh, 2nd Workshop on Magneto- Plasma-Aerodynamics in Aerospace Applications, Moscow, April 2000.
33. “Strong Action of Magnetic and Electric Fields on Inlet Shock Configuration in Diffuser,” S. V. Bobashev, A. B. Erofeev, T. A. Lapushkina, S. A. Poniaev, V. A. Sakharov, R. V. Vasil’eva, and D. M. Van Wie, AIAA-2001-2878, June 2001.
34. “Experiments on MHD Control of Attached Shocks in Diffuser,” S. V. Bobashev, A. V. Erofeev, T. A. Lapushkina, S. A. Poniaev, R. V. Vasil’eva, and D. M. Van Wie, AIAA-2003-0169, Jan. 2003.
35. “Numerical Investigation of Non-Equilibrium MGD Flows in Supersonic Intakes,” Yu. P. Golovachov, Yu. A. Kurakin, A. A. Schmidt, and D. M. Van Wie, AIAA-2001-2883, June 2001.
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37. “Flightweight Magnets for Space Applications Using Carbon Nanotubes,” J. N. Chapman, H. J. Schmidt, R. S. Rouff, V. Chandrasekhar, D. A. Dikin, and R. J. Litchford, AIAA-2003-0330, Jan. 2003.
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39. “Surface Microwave Discharge in Supersonic Airflow,” V. M. Shibkov, V. A. Chernikov, A. P. Ershov, S. A. Dvinin, Ch. N. Raffoul, L. V. Shibkova, I. B. Timofeev, D. M. Van Wie, D. A. Vinogradov, and A. V. Voskanyan, AIAA-2001-3087, June 2001.
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41. “Hypersonic/Supersonic Flow Control by Electro-Discharge Plasma Application,” S. B. Leonov and V. A. Bityurin, AIAA-2002-5209, Oct. 2002.
42. “Influence of Surface Electrical Discharge on Friction of Plate in Subsonic and Transonic Airflow,” S. Leonov, V. Bityurin, N. Savischenko, A. Yuriev, and V. Gromov, AIAA-2001-0640, January 2001.
43. “Effect of Surface Plasma Discharges on Boundary Layers at Mach 5,” Kimmel, R. L., Hayes, J. R., Menart, J. A., Shang, J. AIAA 2004-0509, January 2004.
44. “SparkJet Actuators for Flow Control,” K. Grossman, B. Cybyk, and D. Van Wie, AIAA-2003-0057, Jan. 2003. 45. “Sliding Discharge Applications in Aerodynamics,” V. Bychkov, G. Kuz’min, I. Minaev, A. Ruzhadze, and I.
Timofeev,” AIAA-2003-0530, Jan. 2003. 46. “MHD Effect on a Supersonic Weakly Ionized Flow,” P. Palm, R. Meyer, E. Plonjes, A. Bezant, I. V.
Adamovich, and J. W. Rich, AIAA-2002-2246, May 2002. 47. “Effect of Applied Magnetic Field on the Instability of Mach 4.5 Boundary Layer over a Flat Plate,” F. Cheng,
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Fig. 2 Potential plasma-based technologies.
M0
Plasma Injection/Discharges for Blunt
Leading Edges
Friction andHeat TransferDistribution
Boundary Layer
TransitionControl
Off-DesignOperation ControlWith Virtual Cowl
Cowl Lip Drag and
Shock-ShockInteraction
Control
Shock-Wave/ Boundary-Layer
InteractionControl
MHD Power Extraction and
Flow ProfileControl
Plasma Shortened
Isolator Operation
Inlet ShockPosition Control
Fig. 2 Potential plasma-based technologies.
M0
Plasma Injection/Discharges for Blunt
Leading Edges
Friction andHeat TransferDistribution
Boundary Layer
TransitionControl
Off-DesignOperation ControlWith Virtual Cowl
Cowl Lip Drag and
Shock-ShockInteraction
Control
Shock-Wave/ Boundary-Layer
InteractionControl
MHD Power Extraction and
Flow ProfileControl
Plasma Shortened
Isolator Operation
Inlet ShockPosition Control
Fig. 1 Inlet design challenges.
M0
Blunt LeadingEdge Effects
Friction andHeat TransferDistribution
Boundary Layer Transition
Off-DesignOperation
Cowl Lip Drag and
Shock-ShockInteraction
Shock-Wave/ Boundary-Layer
Interactions
Exhaust Flow
Profiles
Isolator Operation
Fig. 1 Inlet design challenges.
M0
Blunt LeadingEdge Effects
Friction andHeat TransferDistribution
Boundary Layer Transition
Off-DesignOperation
Cowl Lip Drag and
Shock-ShockInteraction
Shock-Wave/ Boundary-Layer
Interactions
Exhaust Flow
Profiles
Isolator Operation
a) AJAX Energy Bypass Engine
b) Aft energy extraction
c) MHD auxiliary Power System
d) External MHD Power Generation
Fig. 3 MHD-based energy management engines.
a) AJAX Energy Bypass Engine a) AJAX Energy Bypass Engine
b) Aft energy extraction b) Aft energy extraction
c) MHD auxiliary Power System c) MHD auxiliary Power System
d) External MHD Power Generation d) External MHD Power Generation
Fig. 3 MHD-based energy management engines.
0
500
1000
1500
2000
2500
0 5 10 15 20 25Freestream Mach Number
Inle
t Tem
pera
ture
(K)
0.24 atm
0.47 atm0.71 atm0.96 atm
Fig. 4 Temperature of inviscid flow entering optimum4-shock inlet.
0
500
1000
1500
2000
2500
0 5 10 15 20 25Freestream Mach Number
Inle
t Tem
pera
ture
(K)
0.24 atm
0.47 atm0.71 atm0.96 atm
Fig. 4 Temperature of inviscid flow entering optimum4-shock inlet.
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a) Over-Sped InletShock Control
b) Under-Sped Inlet Shock Control
c) Virtual Cowl for Under-Sped Operation d) Isolator Control
Fig. 5 Large-scale inlet for control techniques.
a) Over-Sped InletShock Control
b) Under-Sped Inlet Shock Control
c) Virtual Cowl for Under-Sped Operation d) Isolator Control
a) Over-Sped InletShock Control
b) Under-Sped Inlet Shock Control
c) Virtual Cowl for Under-Sped Operation d) Isolator Control
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