UNCLASSIFIED//FOP OFFICIAL USE OHL¥ 21 November 2010 ICO D: 20 July 2010 DIA-08-1011-006 Defense Intelligence Reference Document D efense Futures MHD Air B reathing Propulsion and Power for Aerospace Applications U NCLASSI FI ED//TOR OFFICIAL USE ONLT UNCLASSIFIED//FnB n»Kifth UGC OHb¥ MHD Air B reathing Propulsion and Power for Aerospace Applications The Defense Intelligence Reference Document provides nonsubstantive but authoritative reference inform ation related to intelligence topics or m ethodologies. Prepared by: Technology Warning Division (DWO-4) Defense Warning Office Directorate for Analysis Defense Intelligence Agency Authors: AAP Person 86, AAP Person 87 (U) COPYRIGHT WARNING: Further dissemination of the photographs in this publication is not authorized. This product is one of a series of advanced technology reports produced in FY 2010 under the Defense Intelligence Agency, Defense Warning Office's Advanced Aerospace Weapons System Applications (AAWSA) Program, Comments or questions pertaining to this document should be addressed to[AAP Person 1 | AAWSA Program Manager, Defense Intelligence Agency, ATTN: JUIAF - DI/DWO-3, Bldg 6000, Washington D.C. 20340-5100 ii UNCLASSIFIED//fen OrriCIAL 066 ONLY UNCLASSIFIED//FOR OFFICIAL USE ONLY Contents SUMMARY...................................................................................................iv Chapter 1: CONCEPT OVERVIEW................................................................1 Weakly Ionized Plasmas for Propulsion Applications........................2 Electric Propulsion Systems...............................................................5 Chapter 2: AERONAUTICAL APPLICATIONS............................................11 B asic Principles of Magnetohydrodynamics and Requirements for MHD Performance.............................................................................11 Nonequitibrium MHD in Cold Air Flows.............................................13 The Ajax Concept: MHD B ypass........................................................15 The Reverse Energy B ypass...............................................................18 MHD Applications to Reentry and Near-Orbital Flight.....................20 Chapter 3: SPACE APPLICATIONS...........................................................23 Chapter 4: SUMMARY AND PREDICTIONS...............................................25 Chapter 5: ENDNOTES.............................................................................26 Figures Figure 1. Electrothermal Arcjet Thruster....................................................7 Figure 2. Electrostatic Gridded Ion Thrusters............................................8 Figure 3. Field Orientation for Hall Field Systems and P5 Hall Thruster.............................................................................................8 Figure 4. Electromagnetic Accelerator Field Configuration and Self­ Field Electromagnetic Spacecraft Thrusters.....................................9 Figure 5. Air-B reathing MHD Engine.........................................................10 Figure 6. MHD Control of Scramjet Inlet Using E-B eam Ionization........15 Figure 7. Schematic of Ajax Hypersonic Vehicle Concept........................16 Figure 8. The Reverse Energy B ypass Concept.........................................19 Figure 9. Schematic of the Virtual Cowl Concept.....................................20 Figure 10 . Reentry Vehicle with Surface-Integrated MHD Device and Plasma-Enabled Virtual Streamlining and L/D Increase....................21 Figure 11. Electrothermal Arcjet Thruster on Satellite............................23 Figure 12. SP-10 0 Space Nuclear Power System.....................................24 Figure 13. Nuclear Electric Propulsion (NEP) Concept Vehicles.............24 tii UNCLASSIFIED//FOR OFFICIAL USE ONLY UNCLASSIFIED//TOR. OFFICIAL USE ONL¥ MHD Air-B reathing Propulsion and Power for Aerospace Applications Summary The paper reviews novel propulsion concepts utilizing plasmas (ionized gases) and magnetohydrodynamics (MHD). These concepts are shown to be attractive due to their potential to achieve propulsion and aerodynamic performance far beyond current conventional technologies. However, significant difficulties impede the development and application of these technologies; these include weight, complexity, higher power, and the need for complex and energy-consuming artificial ionization in "cold" air (at Mach <12). A well-publicized Ajax concept of MHD energy bypass has been shown to be meaningless below at least Mach 12. In contrast, a new "reverse energy bypass" with Virtual Cowl is potentially practical for air­ breathing hypersonic vehicles. Applications of the Virtual Cowl and other plasma/MHD devices to reentry, global-strike hypersonic gliders, and aeroassisted orbital maneuvering are identified as promising in the near future. The ability of a plasma/MHD system to generate high power onboard and to provide L/D (lift-to-drag ratio) far beyond that possible conventionally makes these applications both feasible and desirable for national defense. However, these applications are also likely to be implemented by nations such as China, Japan, and Russia. The outlook for uses and applications of MHD propulsion could increase dramatically if high-speed (hypersonic) vehicles begin to carry powerful onboard electricity sources, such as nuclear (fission or fusion) reactors. For spacecraft, the current trend of replacing chemical rockets with electric propulsion systems will continue and is likely to become the standard. Electric systems can provide a much wider range of operation (e.g., low-thrust fine positioning/pointing, more frequent or nontraditional maneuvers, and longer times on station) than chemical systems can. UNCLASSIFIED//FnP PFFTIHAL UDE ONET UNCLASSIFIED//rOR OrriCIAL USE ON LT Chapter 1: Concept Overview A flight vehicle's speed and altitude limit its available propulsion options. T raditional air-breathing systems (propeller, turbofan, and turbojet) are typically limited to altitudes below 80,000 feet. T he existing and planned high-altitude vehicles utilize either high-speed propulsion with ramjet and scramjet engines or slow-speed systems with large propellers. Chemical rockets can operate at all altitudes but have limited burn times and require both fuel and oxidizer to be carried onboard. High-speed air-breathing propulsion, based on ram/scramjet engines, have well- known difficulties: external and internal flow compression and shock control; shock-shock and shock-boundary layer interactions in the propulsion flowpath; mixing, ignition, and flameholding in the combustor; incomplete combustion and chemical energy release; and very high temperatures and wall heat fluxes in the combustor. T here are limits to what can be done about these problems with conventional technologies, which is why the use of plasma (ionized gas) with or without electric and magnetic fields can offer additional opportunities for control and propulsion enhancement. O nboard generation and storage of electric power is one of the main problems encountered with respect to high-altitude, high-speed flight. Hypersonic vehicles, both air-breathing and unpowered reentry "gliders," have no rotating turbomachinery to which an electrical generator could be connected. An attractive power option can be offered by magnetohydrodynamic (M HD) devices. For example, placing an M HD generator immediately downstream of a scramjet combustor can, given the high velocities and temperature of the flow and with metallic additives to the fuel, provide high power (from tens of kW to several M W) with no moving parts. For reentry vehicles, both external (i.e., surface- integrated) and internal-duct M HD generators can generate high power also without moving parts. Employing additional equipment like electrical generators imposes a weight penalty that must be optimized with vehicle performance. T he use of electric and magnetic systems can open new potential areas for aerospace propulsion. While chemical energy sources are limited by the energy available for particular reactions and are limited to operating conditions that are conducive to combustion, electromagnetic energy can be added to the flow over a much wider range of operating conditions. For example, at very high altitudes (>150 kft), it is difficult to get reliable combustion in hypersonic air-breathing engines. In an electrothermal system, the combustor would be replaced with an electrical heating source that can easily and reliably add enthalpy to the flow even at low pressure. T he flow can also be accelerated by manipulating body forces (electric and magnetic) on charged particles (ion and electrons) within the flow. O utside the atmosphere, we note that operation in space almost always requires rocket propulsion whether it be chemical, electric, or nuclear. T he exceptions would be sails and tethers. Spacecraft are rapidly transitioning from chemical rockets to electric (plasma) for most space-based operations.1 T he higher specific impulse (Isp) available for electric systems (2 to 100 times that of chemical) has a dramatic impact on the vehicle design and operation. Although 1 UNCLASSIFIED)Wk UI I 1L1M L USE ONLY UNCLASSIFIED//6Qn QrrTGTAI USE CNET electric (plasma) thrusters have been around for decades, their use in space was limited by the electric power available onboard the spacecraft.2 T he advent of high-power solar arrays has made systems from a few kW to tens of kW practical. Chemical systems will probably always be the primary choice for getting vehicles into space. T he thrust levels for electric systems are too low to be practical for that purpose. Chemical and electric (or electromagnetic) propulsion systems have intrinsic differences. For example, chemical propulsion is "energy limited" because the chemical reactants have a finite amount of energy per unit mass (i.e., their enthalpy of combustion or reaction), which ultimately limits their achievable exhaust velocity. However, because the propellants are their own energy source, the rate at which energy is supplied to the propellant (which is ultimately limited by the reaction kinetics) is independent of the mass of propellant, so very high powers and thrust levels can be achieved. By contrast, electric propulsion systems are typically not energy limited; an arbitrarily large amount of energy can be delivered (from the external solar, nuclear or chemical power system) to a given mass of propellant so that the exhaust velocity can be an order-of- magnitude larger than that of a chemical system. Instead, electric propulsion systems are "power limited" because the rate at which energy from the external source is supplied to the propellant is proportional to the mass of the power system. T his has the result of limiting the thrust of the electric propulsion system for a given vehicle mass. Because of this, electric propulsion vehicles are typically low thrust-to-weight (T /W) ratio (i.e., low acceleration) vehicles. WEAKLY IONIZED PLASMAS FOR PROPULSION APPLICATIONS T his review is devoted to a group of emerging technologies centering on weakly ionized plasmas for propulsion and power.3 Charged particles (ions and electrons) must be present in the flow so that it can interact with applied electric and magnetic fields. Space thrusters operate at very low pressures (< 100 mT orr or < about 2 psi) with a significant fraction of the working fluid/gas being partially ionized (from a few percent to nearly 100 percent). In contrast, air­ breathing systems operate at much higher pressures and have low ionization fractions. T he ionization fraction of concern (i.e., the fraction of gas molecules that are ionized) ranges from as low as IO -8 to IO 2, hence the term "weakly ionized." T he gas pressure in the plasmas can take almost any value. In applications to high-altitude flight, the static pressure is on the order of 10-100 T orr, whereas combustion applications demand near-atmospheric (~760 T orr) or above-atmospheric pressures. T he temperature of the gas can be near-ambient in low-pressure glow discharges, rising to 5,000-10,000K in arc or high-pressure microwave discharges, or even 20,000-30,000K in laser-generated sparks. T he plasmas can be generated by electric or electromagnetic fields, from DC to RF, short pulses, microwaves, and optical (laser) beams, or by various combinations of the above. In general, low pressure plasmas tend to be uniform (diffuse) and nonequilibrium. T he temperature of electrons and internal molecular modes can be very high, while the gas as a whole stays relatively cold. As the pressure and power loading increase, plasmas tend to become hotter, getting closer to thermal equilibrium, and also break into channels (streamers and arcs). T he reality, however, is more complex. In some devices, such as dielectric barrier discharges, nonequilibrium plasmas are generated even at atmospheric pressure, 2 UNCLASSIFIED/FOR OmCIAL UOE ONEY UNCLASSIFIED//rOR OFFIOIAL USE CM LT and in devices such as the gliding arc, the plasma evolves from near-equilibrium to highly nonequilibrium during each of the periodically repeating cycles. In shock and boundary layers during reentry, the plasma is near thermal equilibrium while being diffuse. T he primary reason for this behavior is that the ionization in those shock and boundary layers exists without any electric field and thus is not subject to arcing instabilities. Plasma Features What features or properties make weakly ionized plasmas interesting for propulsion and aerodynamic applications? T he most obvious feature is heating—a consequence of Joule dissipation in an electrically conducting medium placed in an electric field. As a heating element, plasma has important advantages compared with conventional heaters. For example, even a surface electric discharge can effectively heat the gas flow much farther from the wall than a wall-imbedded conventional heater would. M icrowave and laser beams can create plasmas and heat the gas even far from any surfaces, and the volume and shape of the heated region can, in principle, be adjusted. Since heated regions can significantly alter the flow by making the gas flow mostly around them, plasmas can form switchable, controllable, and tunable virtual bodies or surfaces. Such virtual surfaces can be deployed on demand for drag reduction, aerodynamic control (when applied asymmetrically), and optimization of engine inlet performance, to name a few. It is the localized and transient deployment of plasma virtual surfaces that results in the most interesting and complex interactions with gas flows while saving energy compared with large-volume, steady-state plasma utilization, and thus is especially promising for applications. Another useful application of plasma heating is ignition. T his may seem trivial; after all, spark plugs in conventional internal combustion engines are well- developed thermal plasma devices. However, thermal plasma ignition for scramjet engines is not that simple, since the ignition system would have to prevent the plasma from being easily blown away by the supersonic flow, and even if this problem is resolved, if not properly (and quite ingeniously) designed, the igniter would cause an unacceptably strong perturbation to the flow and loss of the stagnation pressure and would require extremely high power. As an example, plasma igniters based on subcritical microwave discharges are quite sophisticated. Besides heating, the presence of charged particles is another obvious, and very important, feature of plasmas. Charged particles can be acted upon by electric and magnetic fields, and this action can be transferred to the bulk gas by ion­ molecule collisions. T hus, magnetohydrodynamic (M HD) and electrohydrodynamic (EHD) interactions can be utilized to exert forces and to decelerate or accelerate the gas in both inviscid core flows and viscous boundary layers. T he magnitude of such interactions depends on the ionization fraction and the magnetic or electric field strength. MHD Interactions T he ionization fraction can be quite high in shock and boundary layers at very high M ach numbers (such as those in reentry flight), or just downstream of 3 UNCLASSIFIED/FOA OFFICIAL USE ONLY UNCLASSIFIED//ron OrriCIAL UDE ONfe¥ ram/scramjet combustors if alkali vapor is added to the gas. In those regions, M HD interactions can be promising for electric power generation or acceleration of the flow, as well as for flow control. However, at M ach numbers below about 12 (and excluding the combustor or the region just downstream of it), the air is too cold for a significant thermal ionization even with alkali seeding. T he required level of ionization then has to be created and sustained by nonequilibrium (nonthermal) means and is associated with a very substantial power budget and additional heating. T herefore, the efficiency of ionization (which can vary by orders of magnitude depending on the particular means of ionization) is of first- order significance for the entire operation and efficiency of the device. Energy used to ionize and excite the gas molecules can be considered as loss in the system since this energy is rarely recovered in the form of directed kinetic or thrust energy. Note that in this regard, ionization by high-energy electron beams or by repetitive high-voltage nanosecond pulses are promising as the most energy-efficient means of nonequilibrium ionization.4'5'6'7 Even with the most efficient ionization techniques, the power budget and additional heating associated with the ionizer normally limit the achievable level of ionization. T o have a substantial M HD effect,8-9 one has to either use a very strong magnetic field (which is associated with some practical issues) or use the M HD interaction in a localized and transient regime (e.g., for boundary layer control). As for EHD interaction,10'11 it relies upon non-neutrality of the plasma and an electric field to impart momentum to the gas. Although EHD (or "ion wind") phenomena have been known for many years, the last several years saw a surge of new interest to this type of interaction. T his new boom is due to the asymmetric dielectric barrier discharge (DBD)—a remarkably simple device that has been demonstrated to be very effective in delaying and controlling flow separation and perhaps even laminar-turbulent transition. Although details of the physics of DBD plasma actuators are still incompletely understood, the simplicity of these devices, their low power consumption, and the striking effectiveness in separation control bring these systems to the top of the list of plasma aerodynamics and plasma-assisted propulsion technologies that have near-term application prospects. Combustion Another area where nonequilibrium (nonthermal) weakly ionized plasmas are very promising is plasma-assisted combustion. Although heating induced by plasmas can ignite combustible mixtures, as mentioned above, it is the presence of "hot" electrons in a cold gas that makes nonequilibrium plasmas quite interesting for promoting chemical processes such as combustion. Electron - impact dissociation, excitation, and ionization of molecules can generate chemically active species such as radicals and excited atoms and molecules, and those species can initiate or accelerate chemical reactions that would otherwise be nonexistent or slow at low temperature. A number of novel techniques, including (but not limited to) high-voltage nanosecond pulses and the so-called "gliding arc" have been shown to be quite effective in plasma-assisted combustion. Investigation of detailed mechanisms (often quite complex and 4 UNCLASSIFIED/£CML4SHfiMM6MN» UNCLASSIFIED//FOR OFFICIAL USE ONLY nontrivial) of the coupled physical and chemical processes in those plasmas can potentially lead to their better understanding and help them become practical. ELECTRIC PROPULSION SYSTEMS Electric propulsion thrusters can be divided into three categories: electrothermal, electrostatic, and electromagnetic. First, electrothermal thrusters use electric energy to directly heat the propellant and add enthalpy. T he heated gas is then accelerated using a conventional converging-diverging gas-dynamic nozzle. Second, electrostatic thrusters use applied static electric fields to accelerate propellant ions via body forces. T hird, electromagnetic thrusters use electromagnetic body forces (ExB) to accelerate a plasma (positive and negative charges). An electric propulsion system consists of a power source (e.g., solar or nuclear), power conditioning electronics, engine/thruster (including inlet for air­ breathing systems), and fuel/propellant storage and feed subsystem. Energy can be obtained from sunlight, a nuclear reactor, or chemical sources. In the case of solar electric propulsion (SEP), solar photons are converted into electricity by solar cells. T he energy could also be beamed to the vehicle using laser or microwave sources. Beaming the power allows for higher power densities but with the added complications of needing a power station and a means of getting the power to the vehicle (direct illumination or via a relay system). In nuclear electric propulsion (NEP), thermal energy from the nuclear reactor is converted into electricity by either a static or dynamic thermal-to- electric power conversion system. Static systems have the advantage of no moving parts for high reliability, but they have low efficiency; dynamic systems have moving parts (e.g., turbines and generators) and do not scale well for small systems, but they do have higher efficiency. O ther onboard energy storage systems such as high-density capacitors, flywheels, or fuel cells could be used. Power conditioning systems are required to convert the power system voltage to the form required by the electric thruster. For example, an SEP power system produces low-voltage DC (typically ~100V); this would need to be converted (via transformers, etc.) to kilovolt levels for use in an ion thruster. T he power­ conditioning system is often referred to as the power processing unit (PPU); this is, in turn, part of the vehicle's overall power management and distribution (PM AD) subsystem. Various combinations of thruster and propellant are possible, depending on the specific application. T he propellant or working fluid can be either stored on board and used in a rocket mode or collected from the atmosphere in an air-breathing mode. T he natural system-level trade between these propellant methods is fuel/propellant mass versus power system mass. Air-breathing systems require less propellant mass but require higher energies to perform similar missions. Although rockets will operate in a space or air environment, their thrust durations are limited by the amount of propellant that can be carried. Key performance parameters determine the relative strengths and weaknesses of different propulsion technologies. T he fuel/propellant efficiency is characterized by the specific impulse (Isp) for rockets and by the thrust-specific- fuel-consumption (T SFC) for air-breathing systems. It is a measure of how much 5 UNCLASSIFIED/EQR OFFICIAL USE OM E* UNCLASSIFIED/ XEQUEHGlAb-WGMNteV thrust (Fth) is produced from each mass unit of propellant. T he thrust efficiency (q) is a measure of how much of the power/energy that is available results in directed kinetic energy (thrust power) in the flow. T he thrust density is a measure of how much thrust is produced per unit cross-sectional area, Ac (area perpendicular to the flow direction). Although the thrust density is a packaging issue for spacecraft, it is critical for air vehicles where drag is present. T he thrust-to-power measures the acceleration efficiency. T raditionally a trade exists between fuel/propellant efficiency and thrust-to-power (speed vs. economy). T he final parameter is the specific mass or weight (mass)-to-power ratio. In most cases the efficiencies improve with the size of the system (economy of scale). M easures of performance are fundamentally different between air-breathing and rocket systems due to the inlet on the air-breathing system. Rocket Air-Breathing = m e F ™ = m e Ue + Ae(Pe ' Pan* ) = ^ u^ _ Thrust propellant weight flow rate _ Fth _ Ucq m e go 8° FW = maUe-m«U + A=(Fe - Pant ) = W « [(1 + f k - «] + Ae (Pe - Pm* ) m e = m f + m a f = m f ! m a _ Thrust _ Fth fuel weight flow rate Pja 1/2 m e ueq — 1/2 Flh ISP g0 fuel mass flow rate _ tn? Thrust Fth Pj« =1/2 m. u^ " m a «' « = Jj£L - F o l1spJ> o . _ I/2 ^f ueq P 2 P Pelect elect elect P * elect 2 2 ^Ueq -m ou Above, m a is the mass flow rate of the air, m f is the mass flow rate of the fuel, m e is the exit flow rate, g0 is the acceleration of gravity (reference point, Earth), ue is the nozzle exit velocity, u is the flight/vehicle velocity, ueq is the equivalent exit velocity, Ae is the nozzle exit cross-sectional area, Pe is the nozzle exit 6 UNCLASSIFIED/FOA OFFICIAL UOE ONLY UNCLASSIFIED//TOW OFFICIAL UOC ONLY pressure, Pamb is the ambient/environmental pressure, Pjet is the jet-kinetic power (thrust power) produced by the engine, and Peiect is the electrical (or other external source) power supplied to the propulsion system. For air-breathing systems the mass flow rate of the fuel is very small relative to the mass flow rate of the air. Note that an air-breathing system can never fly faster than its exhaust velocity. Rockets, because they carry an onboard oxidizer, do not have this restriction and, therefore, have no flight-speed limits. T he jet power for high-speed air­ breathing engines is larger than the equivalent jet power for a rocket due to the inlet and is the difference of two large numbers. Large power levels are required for aircraft and launch vehicles. For example, an SR-71 cruising at M ach 3.2 produces a thrust of 24,700 Ibf (110 kN) and a jet power of 104 M W. Climb and maneuver thrust is much higher. Similarly, an RL10 rocket engine produces 15,000 Ibf (66.7 kN) thrust at an ISp of 433 seconds, and has a jet power of 142 M W. By comparison, a Nimitz class nuclear aircraft carrier propulsion system is 194 M W, and the Hoover dam produces about 2000 M W. T herefore, any electric system replacing these applications must be capable of processing a lot of power. Spacecraft propulsion systems are typically hundreds of watts to tens of kW. T his, in addition to powerplant weight issues, is a primary reason why electric propulsion systems are currently being used on spacecraft and not on aircraft. Historically, electric thrusters for spacecraft were available for flight decades before the power systems.12 Figure 1. Electrothermal Arcjet Thruster. Photograph of a 30-kW arcjet thruster being tested at the Jet Propulsion Laboratory.13 7 UNCLASSIFIED/rOR OFFICIAL USE ONLY UNCLASSIFIED//rne AEriGiAI UOC OH LT Electrothermal thrusters use electric energy to heat the propellant and add additional enthalpy. T his can be done with simple resistive heating or by passing the propellant gas through an arc plasma discharge. T he plasma can be generated through a high-current discharge or by absorption of microwaves. T he hot pressurized gas is then accelerated out of the thruster using a conventional converging-diverging gas-dynamic nozzle. An example of an electric arc heated thruster or arcjet thruster is shown in Figure 1. Electrostatic thrusters use an applied static electric field to accelerate propellant ions. Strong electric fields are created in the engine which then accelerate the (positive) ions to high velo­ Figure 2. Electrostatic Gridded Ion Thrusters. Photo­ graph of a gridded 30 cm diam­ eter ion thruster being tested at the Jet Propulsion Laboratory.14 cities. T he accelerating field can be applied using physical grids such as those used in ion engines or using "virtual grids" generated by an applied magnetic field that traps the electrons as is done in Hall-effect thrusters. A photograph of the NASA ion engine used on the Deep Space O ne spacecraft is shown in Figure 2. While gridded electrostatic thrusters like ion thrusters are cap­ able of very high Isp (1,000 to >20,000 seconds) values they have very low thrust densities (1-5 N/m2) due to the space-charge current limit in the accelerator system. Hall-effect thrusters do not have this space-charge limit but also have thrust density limits due to the annular geometry (tens of N/m2). T ypical power levels are from watts to 50 kW. In the Hall field orientation, the electric field causes electrons to flow upstream and the ions to drift toward the exhaust as shown in Figure 3.15 T he electrons and ions transfer equal and opposite amounts of momentum to the air, resulting in zero thrust when no magnetic field is present. However, with the application of a transverse magnetic field, the forward flow of electrons is slowed while the aft flow of ions is nearly unaffected. Consequently, there is a net momentum transfer resulting in thrust on the vehicle. Hall Field Orientation Figure 3. Field Orientation for Hall Field Systems and P5 Hall Thruster. Left: Figure shows the Hall field orientation. Right: P5 Hall Effect thruster being tested at the University of M ichigan. 8 UNCLASSIFIED/rOR OFFICIAL USE ONLY UNCLASSIFIED//rQR OmCIAL UOC OHL-* Electromagnetic thrusters use electromagnetic body forces (ExB ) of Lorentz force to accelerate a propellant plasma as shown in Figure 4.16-17 T he electric field is applied using electrodes within the thruster. T he Lorentz or ExB force accelerates both positively charged ions and negatively charged electrons or negative ions in the same direction. T he magnetic field can either be applied externally (applied field thruster) or generated by a very high current (typically thousands of amps) plasma discharge (self-field thruster). T he current also serves to ionize the propellant. It is the interaction of the electric and magnetic fields that pushes the plasma out of the thruster at high velocity via the Lorentz force that acts mutually perpendicular to the electric and magnetic fields. Spacecraft electromagnetic thrusters are capable of processing much higher power densities (50 kW to tens of M W) and much higher thrust densities than electrostatic thrusters (hundreds to thousands of N/m2). Figure 4. Electromagnetic Accelerator Field Configuration and Self-Field Electromagnetic Spacecraft Thrusters. Left: Illustration of the electric and magnetic field configuration. Center: Photograph of a M W-class pulsed magnetoplasmadynamic (M PD) thruster being tested at Princeton University. Right: Photograph of a 50-kW steady-state M PD thruster being tested at Princeton University. An example of an air-breathing electromagnetic accelerator system for high­ speed and high-altitude flight is shown in Figure 5. Air enters through the inlet on the left. T he center section both accelerates the flow via electromagnetic body forces and heats the gas through O hmic heating. T he gas is then further accelerated using a diverging gas-dynamic nozzle. 9 UNCLASSIFIED/mB QM JCWM USE ONL¥ UNCLASSIFIED//FQR OrriCIAL UOE OH EV Figure 5. Air-B reathing MHD Engine. An illustration of a conceptual high-speed air­ breathing M HD engine and (insert) a photograph of the operating proof-of-concept experiment being investigated by Lockheed M artin Aeronautics. We note here that in its first-ever list of top 10 emerging aerospace technologies, released in 2009, the American Institute of Aeronautics and Astronautics (AIAA) included two plasma technologies: plasma actuators for active flow control and plasma-based advanced space propulsion technologies. 10 UNCLASSIFIED/rOR OFFICIAL USE ON FT UNCLASSIFIED//TOR OTTICIAL USE ONL'f Chapter 2: Aeronautical Applications - Concepts and System Issues In this section we will review the following issues and concepts: • Basic principles and problems of M HD propulsion, power generation, and flow control. • M HD inlet control. • M HD power generation in scramjet flowpath. • Plasma-generated virtual surfaces for drag reduction, steering, and virtual cowl. • M HD energy bypass: the Ajax concept. • T he reverse energy bypass concept. • M HD power generation and aerodynamic control for reentry vehicles. B ASIC PRINCIPLES OF MAGNETOHYDRODYNAMICS AND REQUIREMENTS FOR MHD PERFORMANCE T he basic principles of magnetohydrodynamics (M HD) are understood very well. When an electrically conducting fluid crosses magnetic field lines, an electromotive force (Faraday e.m.f,, equal to the product of flow velocity u and the strength of magnetic field B , uB, multiplied by the channel width) is induced across the fluid and the B field. If then a pair of electrodes is positioned on either side of the fluid flow and connected via a ballast resistor on the outside, an electric current will be induced in the circuit, and power will be generated on the external load. T his electric power will represent partial conversion of the flow enthalpy (consisting of thermal and kinetic energy of the flow) into electricity. At the same time, the current flowing through the finite-conductivity fluid will produce Joule heating3 of the fluid that will increase both static temperature and entropy of the fluid. T he ratio of the extracted electrical power to the Joule dissipation rate is determined by the ratio of the load resistance to the sum of load and fluid resistances; this ratio is called the "load factor," k, 010-100 T orr) are much higher than those in typical glow discharges (1 T orr or less) resulting in much higher power required to sustain plasmas and to severe problems with arcing instabilities. • T he ionization fraction needed for a good electrical conductivity and acceptable M HD performance is much higher than that required for a fluorescent light, again resulting in high power budget and overheating. For cold nonequilibrium plasmas, the power budget is determined by the average energy cost (usually expressed in eV), W i, of ionization (i.e., of producing an electron-ion pair), and the rate at which the electron-ion pairs must be generated in order to compensate for electron losses in recombination, attachment, and other processes. T he recombination is the dominant loss mechanism at reasonably high electron densities, and its rate is proportional to the product of electron and ion number densities. Since in quasineutral plasmas the number densities of electrons and ions are close to each other, the recombination rate (per unit volume) is equal to kdrne2, where kdr is the dissociative recombination rate coefficient and ne is the electron number density. Note that the characteristic plasma decay time due to recombination is almost always very short, typically ~l-10 microseconds, so that the flow moves only a very short (~1 cm) distance during the decay time. T his is why schemes with pre-ionization upstream of the M HD region with no ionization in the M HD region itself are not viable; the ionization must be done continuously throughout the M HD region. 13 UNCLASSIFIED/EQR OFFICIAL UDE ONLY UNCLASSIFIED//TOR UHP1L1M L USE ONL¥ T he average energy cost, M A, of ionization varies greatly depending on the ionization method. For example, in conventional glow-like discharges of large volume at moderate or high pressure, the ionization cost is ~10,000 eV (i.e., three orders of magnitude higher than the minimum ionization energy [10-15 eV]). T his is due to the low average electron energy (~1 eV) and to the dominant losses of electron energy in inelastic collisions with air molecules. T his is why a highly efficient ionization technique must be used in order to give cold­ air M HD devices a chance to be viable. High-energy electron beams represent such a technique. Generated in vacuum electron guns and injected into air through either thin foil or a differentially pumped window, energetic (>1-50 keV) electrons produce many more low-energy plasma electrons, so that the average ionization cost is only M A = 34 eV. T his ionization efficiency is theoretically the best. O f course, electron beam systems are quite difficult to work with due to fragile foils or massive differential pumping facilities; X-ray generation is also not helpful for flight applications. But even putting these important practical problems aside, and even with the lowest possible cost per electron, the requirement that a cold-air nonequilibrium M HD device uses significantly less power for ionization than it extracts from (in the generator case) or adds to (in the accelerator case) the flow imposes a severe constraint on the maximum level of ionization and conductivity. Calculations show that the maximum ionization fraction is on the order of 10'6 and the maximum conductivity is on the order of 1 mho/m. With this low conductivity, substantial (S~0.1 or higher) M HD interaction parameters can only be reached with magnetic fields higher than several T esla (i.e., 10-20 T esla). T he weight and volume of a magnet then makes such flight devices quite impractical, unless a breakthrough in magnet and materials technologies occurs resulting in ultralightweight magnets with B~10 T esla. As an example of potential use of nonequilibrium cold-air M HD devices with ionization by e-beams, we note the studies of M HD scramjet inlet control performed by one of the authors of this survey and his Princeton University colleagues. T hese theoretical/computational studies showed that indeed, with proper optimization, M HD interaction at the compression ramp upstream of the scramjet inlet can restore the shock-on-lip (SO L) condition at M ach numbers higher than the design M ach number for a given fixed-geometry inlet (Figure 6). During the M HD operation, the generated electrical power would be enough for ionizing e-beams, with a significant percentage of the power left to be stored onboard and used for other purposes. T he advantage of M HD inlet control is that it eliminates the need for a variable-geometry (movable) cowl that would be associated with a large weight and complexity; the disadvantage is that the weight and complexity associated with magnets and e-beam systems may negate the advantages. Systems studies are needed to fully assess the practicality of this M HD inlet control, and results of such studies would strongly depend on the state-of-the-art and future advances in lightweight magnet and e­ beam technology. 18-19-20-21-22 14 UNCLASSIFIED/FOB PFrilSIAL USE ONET UNCLASSIFIED//fOW OFFICIAL USE ONL¥ Figure 6. MHD Control of Scramjet Inlet Using E-B eam Ionization. T he retarding ampere force restores the shock-on-lip condition at M ach numbers higher than the design value. THE AJAX CONCEPT: MHD B YPASS An M HD-assisted propulsion concept that has attracted perhaps the most attention over the last decade or two is known as the Ajax, or Ayaks. T he concept, illustrated in Figure 7, originated in the 1980s at Leninetz Scientific & Production Enterprise (now Leninetz Holding Co.) in Leningrad, USSR (now St. Petersburg, Russia). A good overview of the concept and its present status and problems can be found in the recent article23 included in the Special Section of the Journal of Propulsion and Power devoted to weakly ionized gases for propulsion enhancement, with one of the authors of this survey serving as special guest editor. 15 UNCLASSiFiED/rnn QrnrT^i ijoe only UNCLASSIFIED//fiQB OFFK^L UGE OHL* -5QQQ2- NozzleKinetic energy of air stream Active thermal protection systems Systems controlling aerodynamic ^^J^’!^.^ characteristics in gas ,ransier sterns and plasma medium Chemical energy of original fuel (70%) - ZZZ^^^ZZZ^^ZZZZZZZ M agnetoplasm ©chemical engine Additional chemical energy obtained | ~ due to heal regeneration (30%) Losses during movement In a continuous medium Figure 7. Schematic of Ajax Hypersonic Vehicle Concept. T he upper and lower pictures represent the same concept, but were published by the authors (A. Kuranov et al.) at different times. T he key idea of the Ajax is that, with propulsion, aerodynamic, and heat protection system for hypersonic vehicles hitting their theoretical and practical limits, the only way beyond these limits is smart management of energy (i.e., taking energy from the surrounding flow and putting that energy where it is needed for propulsion benefits). T he following three major concepts/systems are included in Ajax:24 • Endothermic fuel conversion. A mixture of hydrocarbon fuel (similar to kerosene) and water initially stored on board is used to cool the external surfaces and engine walls, and the heat transferred to the mixture is used in 16 UNCLASSIFIED/ren OFFICIAL USE ONLY [PropelianF^, UNCLASSIFIED//TOW OFFICIAL USE ONLY a thermocatalytic "cracking" process that makes a syngas (i.e., CO - H2 gaseous mixture) from the original kerosene-water liquid mixture. T he syngas made onboard is a much better fuel from the Isp standpoint than liquid hydrocarbons, and with the onboard thermocatalytic conversion there is no need to carry hydrogen from the takeoff. T his part of Ajax is certainly very meaningful and probably viable. • MHD energy bypass. T his is perhaps the most controversial part of Ajax. An M HD generator extracts energy from the airflow upstream of the scramjet combustor; this energy bypasses the combustor and is put back into the flow via M HD accelerator placed downstream of the combustor. We will discuss this concept below. • Plasma for drag reduction. A part of the M HD-generated power can, in principle, be used to generate a plasma in air upstream of the vehicle nose. T his plasma would weaken the bow shock and reduce the wave drag on the hypersonic vehicle. Although there were claims by some Russian groups about 10-15 years ago that weakly ionized plasmas can reduce shock strength via some unknown physical mechanism, extensive research in the United States, Europe, and Russia has conclusively shown that the effects are purely thermal. However, even with purely thermal action, plasma drag reduction can be quite meaningful and useful for high-speed flight (see below). Perhaps the most basic problem with the M HD bypass, as pointed out by D. Riggins,25 is that as a propulsion power cycle, it runs in the direction opposite to that dictated by thermodynamics. Indeed, any thermodynamically correct heat- into-power conversion cycle has work addition (e.g., compression) prior to heat addition (e.g., in the form of combustion), and work extraction follows the heat addition. T his is why air is compressed (work added) upstream of the combustor in ail normal propulsion cycles, whether by compressor in a turbojet or a compression ramp upstream of a scramjet combustor. In this sense, M HD power (work) extraction before air enters the scramjet combustor, followed by M HD power addition after the combustor, constitutes a thermodynamically "wrong" and thus inherently inferior, propulsion system. However, in criticizing the Ajax power cycle and arguing that the Isp of Ajax is always less than that of a system without M HD bypass, D. Riggins26 makes a significant mistake. In his derivations, he assumes that combustion-generated heat addition in the combustor occurs at a gas temperature equal to the stagnation temperature of the flow (i.e., that the flow is fully stagnant in the combustor). T his assumption is in direct contradiction to the very idea of a scramjet, where combustion occurs in supersonic flow. Heat addition in the combustor thus occurs at a static, not stagnation, temperature. It is this fact that at least gives M HD bypass a chance to increase Isp. Indeed, calculations described in the above-referenced paper27 by the Ajax group do result, in some conditions and with careful optimization, in an Isp increase. O ur analysis of their calculations shows that increase in static temperature caused by flow deceleration and Joule dissipation in the M HD generator upstream of the combustor is the reason for higher Isp. Indeed, since the entropy increase in the combustor is equal to Q/T , where Q is the heat added and T is the static temperature at which this heat is added, any increase would lead to lower 17 UNCLASSiFiED/rnn Qrrrri^i use ewer UNCLASsiFiED//ren ornciAL use onb* entropy increase and thus, as can be easily shown, to higher Isp. T he Isp increase due to the increase in combustor temperature is made smaller by negative factors such as irreversibilities due to Joule dissipation in both M HD generator and accelerator and by the thermodynamically "wrong" work extraction before the combustor. T he Isp increase, however, even in optimal cases, is only several percent. Given the crude assumptions in the paper28 (ID flow, no boundary layer and heat losses, uniform plasma, no e-beam energy losses, no losses in electric circuitry), this gain of several percent would turn into a loss of Isp in more realistic analysis. Additionally, the weight and complexity associated with magnet and e-beam systems should be kept in mind. T herefore, one can state with certainty that M HD energy bypass at M ach<12 (where nonequilibrium ionization of air is required) is not a meaningful technology. Where the M HD bypass could be useful is at very high M ach numbers (M ach>12). First, stagnation temperatures at these M ach numbers are high enough for significant thermal ionization with reasonable amount of alkali seed (0.01-1% by volume), thus eliminating the need for a heavy, complex, and entropy-generating nonequilibrium ionization system. Second, at static temperatures (>2,000K) reached in the combustor at these M ach numbers, there is no combustion per se, just dissociation of fuel and air molecules followed by full or partial recombination into other molecules that releases heat into the flow downstream of the combustor in the expansion nozzle. For such a regime, the group at NASA Ames showed29’ 30 through modeling that M HD bypass can indeed increase the IsP. Note, however, that materials and structures, as well as fuel development, are currently such that air-breathing flight at M ach>12 is not realistic. In the future, if air-breathing propulsion at M ach>12 becomes possible in principle, reexamination of M HD bypass benefits and flaws will be warranted, especially if lightweight magnets also become available by that time. THE REVERSE ENERGY B YPASS Returning to M ach<12, one of the authors of this survey, together with his colleagues, has proposed a very different bypass concept, dubbed the reverse energy bypass (Figure 8).31-32 T he energy (in the form of electricity) is extracted from the flow in an M HD generator placed just downstream of the combustor (or collocated with the combustor). T his at least avoids the need for e-beam ionization, since the air mixed with combustion products is sufficiently hot right after the combustor that an acceptable electrical conductivity (on the order of 10 mho/m or higher) can be generated thermally, provided alkali metals are seeded into the fuel and are thus present in the combustor and downstream of it. A part of the electrical energy generated downstream of the combustor could be used upstream of the M HD generator, which is why this is called the reverse energy bypass (the energy being bypassed is moved in the upstream direction). Plasma-assisted combustion (such as ignition, flameholding, and mixing) would benefit from this electrical energy. Plasma heat addition, in steady or transient modes, enabled by this electrical power, would be beneficial for control of shock interaction at the inlet and for drag reduction and/or steering and pitch or yaw control when used in front of the vehicle's nose. 18 UNCLASSIFIED/rUK OFFICIAL USE ONLY UNCLASSIFIED//rng ntHCTftl UGH OM Wf /---------------------------------------------•. Plasma directed energy steering and lift control Magnetically driven surface plasma for boundary layer control and suppression of Plasma enhanced mixing, flame spreading, and ignition control Plasma energy addition for flow/shock control in transient operation Plasma generated virtual cowl lip for air capture increase . in acceleration MHD heat-into-power conversion Figure 8. The Reverse Energy B ypass Concept. T he red arrow symbolically depicts the general direction of energy bypass - upstream. A promising version of the reverse energy bypass concept uses the electric power extracted from the flow in the M HD generator to increase the air mass flow rate through the combustor in off-design conditions. If the inlet is designed for shock-on-lip condition at a certain high M ach number (e.g., M ach 8), then at M ach numbers lower than the design value (e.g., M ach 6) the inlet does not completely capture the compressed flow, which is associated with the so-called spillage drag, effectively reducing the thrust. T his undesirable effect can be prevented by the so-called virtual cowl—a heated region upstream of the cowl lip (Figure 9).33-34 T his heated (e.g., plasma) region deflects the flow and helps with scooping more air into the inlet. M oreover, with proper positioning, only cold (unheated) air is scooped into the inlet, thus avoiding reduction in total pressure and thrust that would have occurred if heated air were scooped into the propulsion system. 19 UNCLASSIFIED/ran OFFICIAL UPC ONLY UNCLASSIFIED//FOR OFFICIAL UOC ONLY Figure 9. Schematic of the Virtual Cowl Concept. Plasma-generated heated region upstream of the cowl lip deflects the flow and scoops more air into the inlet. Analysis of the reverse energy bypass involving M HD generator downstream of the scramjet combustor and an optimized virtual cowl revealed that, although enthalpy extraction and entropy production in the M HD generator substantially reduce the thrust, the combined system with virtual cowl can actually increase thrust by as much as 20-30% .35-36 From the system standpoint, this concept, with all its associated weight and complexity, should be compared with that involving a movable (variable geometry) cowl that is associated with a heavy and complex electrohydraulic system and also requires power. An attractive feature of the reverse bypass concept that might make it a winner is its multifunctional nature. Indeed, M HD power generation downstream of the combustor can be an attractive power­ production option for hypersonic vehicles. Alternatives (batteries, fuel cells, and the like) are not very competitive for generation of large amounts (hundreds of kW to 1-10 M W) of power onboard. T hus, if an M HD generator in the propulsion flowpath is accepted as the power source, its operation in conjunction with a virtual cowl that significantly increases thrust in off-design conditions, and also enables aerodynamic control and maneuvering, would become quite practical. Note also that if another source of high (M W-scale) power, such as a nuclear reactor, is onboard, its use for virtual cowl, drag reduction, and aerodynamic control would be straightforward and would greatly increase performance of the hypersonic vehicle. MHD APPLICATIONS TO REENTRY AND NEAR-ORB ITAL FLIGHT We now turn to M HD application to reentry and near-orbital flight. Due to the high velocities and enthalpies involved, the gas temperature in shock and boundary layers is very high, from several thousand to 10,000-20,000K. At these temperatures, thermal ionization is very substantial, and at the low end of 20 UNCLASSIFIED/TOR OTriCIAL UGE ONh¥ UNCLASSIFIED//FOR OFFICIAL USE ONLY the temperature range, a moderate seeding with NaK (sodium-potassium) mixture would be sufficient to produce an electrical conductivity from >100 mho/m to as high as 1,000-3,000 mho/m. With this level of conductivity, very modest magnetic field B~0.1-0.2 T esla would suffice for a strong M HD performance (Figure 10). M odeling shows that M W-scale power can be generated in these conditions by a surface-in teg rated M HD system from 1 square meter of vehicle surface. Interestingly, calculations show that the additional weight of the system, assuming a 1,000-second mission, is determined mostly by the water required to cool the copper-wire electromagnet and that the additional weight is quite acceptable, increasing the practicality of the system.37- 38-39- 40 ma for virtual amlining and increase Figure 10 . Reentry Vehicle With Surface-Integrated MHD Device and Plasma- Enabled Virtual Streamlining and L/D Increase. O ne good use of such high power would be to create a plasma in front of the vehicle in order to reduce drag (Figure 10). Nonoptimized analysis shows that the "return" (i.e., the drag power saved divided by the power spent on creating the plasma) can easily be as high as 40-50 (i.e., the drag power saved is 40-50 times greater than the power spent on the plasma).41 T here are theoretical and experimental indications that with proper shaping of the plasma region (specifically, making it long and thin), the "return" can be >100. T hus, this "reverse energy bypass" would result in substantial reduction in drag and increase in L/D (lift-to-drag ratio) by tens of percent. T he increase in L/D would directly translate in increased downrange for an unpowered hypersonic global "glider" and in increased cross-range for a de-orbiting space asset. Note that off- 21 UNCLASSIFIED/rOR OFFICIAL USE ONLY UNCLASSIFIED/ /.EOMEHCJAWWe-eNte* axis positioning of the plasma region would create steering or pitch/yaw control moments.42 An attractive application that utilizes this plasma/M HD-enabled increase in L/D is orbit inclination changes for space assets. Even a modest (a few degrees) orbit inclination change requires a very large amount of delta-velocity and energy and thus a very large amount of fuel to be burned. If the space asset dives into the upper atmosphere (to altitudes of 200-300 kft), it can use aerodynamic turning (similar to airplanes), provided the L/D ratio is high enough. Unfortunately, hypersonic L/D, especially in rarefied air at high altitudes, is not much higher than 1. Plasma and M HD technologies hold substantial and realistic promise to achieve hypersonic L/D of 3-10, which would be a game-changer and enable, among other missions, aerodynamically assisted, on-demand orbital inclination changes. We now briefly consider another M HD application: a hybrid chemical/M HD propulsion. T he nozzle exit velocity of chemical systems (air-breathing and rockets) is limited by the chemical energy available from the fuels/propellants and the temperature limits of the system materials. O ne method to increase the exit velocity of the system is to add an M HD accelerator system to the nozzle. T he flow is first accelerated using a conventional gas dynamic converging­ diverging nozzle and then the M HD system further accelerates the supersonic flow in the diverging portion of the nozzle. M any ground-based systems have been developed and tested to accelerate flows using M HD systems. T hese have been primarily either proof-of-concept systems or for hypersonic wind tunnels.43' 44 Systems have been proposed for both small in-space systems45, 46 and for large engines for launch vehicles.47 While this concept has great potential and the accelerator physics are well established, it has several practical limitations. T o be efficient, the energy added to the flow from the M HD system should be on the order of or greater than the energy added by the chemical stage. T his requires power levels that are not available on either type of vehicles. For example, for launch vehicles the jet power levels would be in the hundreds of M W to tens of GW range. T he low ionization fractions in the flows also severely limit the thrust efficiency of the M HD systems to a few percent. T his combined with the large jet powers requires enormous launch-vehicle powers. Similarly for space systems, a better solution would be to use the available power in a more efficient electric thruster. T he large powers also require large masses for the M HD system components for reasonable specific mass (kg/kW). T o be comparable to pure electric systems on spacecraft, the M HD augmentation system specific mass would need to be improved by a factor of 1,000 over state- of-the-art technologies.48 O ne potential solution around the power issue is to beam the power to the vehicle.49 Another serious issue is the magnets needed to provide the 2- to 40-T esla fields required. In many cases, the weight of the magnet and magnet power supply would exceed the vehicle mass using existing technology. T he magnet system mass will need to be reduced by several orders of magnitude to make flight systems practical. 22 UNCLASSIFIED/TOK OFFICIAL USE ONLY UNCLASSIFIED//rOR OFFICIAL BSC ON LT Chapter 3: Space Applications Electric propulsion systems are currently being used for attitude control, positioning, and primary propulsion. T he use of electric propulsion systems on spacecraft was limited by the amount of power available. T he thrusters were developed decades before the power systems. Early applications were the replacement of hydrazine monopropellant thrusters with hydrazine resistojet and arcjet thrusters (increasing Isp from 200 seconds to 300 seconds for resistojet thrusters and to 600 seconds for arcjet thrusters). An example is shown in Figure 11. Figure 11. Electrothermal Arcjet Thruster on Satellite. Lockheed M artin Series 7000 Comsat with Aerojet 1.8-kW Arcjet thrusters (insert photo) for north-south station keeping. T his keeps most spacecraft the same, just changing the thrusters (lower risk and cost). Newer spacecraft are being designed specifically for use with electric thrusters. T hese are primarily gridded ion engines and Hall-effect thrusters operating on xenon propellant. Xenon is a unique noble gas that can be stored with densities close to liquids at pressures above 800 psia. As the available electric power has increased, the transition to all electric spacecraft has increased, as well as the sizes of the electric propulsion systems. Hall thrusters are used for station keeping as well as apogee insertion maneuvers. T his trend will continue for decades to come.50 T he high Isp available from electric systems enables new operation concepts for what a spacecraft can do. T he amount of propellant that can be stored onboard limits the number and types of maneuvers the spacecraft can perform. Electric systems enable enhanced ability to relocate assets, fly nontraditional or non- Kelplerian orbits, and keep spacecraft on station for much longer periods. Although the Isp of electric systems are much higher than chemical systems, the thrust levels are much lower. T his results in much lower spacecraft accelerations and longer repositioning times. T he availability of higher power levels will allow for higher power thrusters to be used and, therefore, the repositioning times to be lower. For a given power, the Isp and thrust can be traded (Peiect = Vz go IsP Fth /q). High-power Hall-effect thrusters are being designed to operate in both a high-thrust (lower Isp) mode for obit insertion and repositioning and high-Isp (low thrust) for propellant-efficient maneuvers and station keeping. T his adds significant flexibility to how the spacecraft is operated and the missions it can perform. Very fine spacecraft positioning and pointing can be accomplished using the low thrust levels associated with some electric systems. For example, field emission electric propulsion (FEEP) and colloid thrusters are capable of thrust levels in the micro-Newton range and can be used to offset small spacecraft perturbations such as solar wind. 23 UNCLASSIFIED/rQn OFFICIAL USE OM ET UNCLASSIFIED//JOB QITTGTAI UOE CHET Figure 12. SP-10 0 Space Nuclear Power System. Illustration of the 100 kWe SP-100 nuclear reactor. Another operational option enabled with cuta w a y vie w of the sp-mo spa ce electric systems is to fly at much lower obits where drag forces would normally cause the spacecraft to reenter in a short period of time. An air-breathing electric thruster can be used for drag make-up without the need for additional propellant.51 Nuclear fission power systems coupled with electric thrusters enable new capabilities. Nuclear space reactors have been flown in space (SNAP-10A by the United States and T O PAZ reactors by the former Soviet Union).52 T he SP-100 system shown in Figure 1253 was being developed in the 1990s as a tug to move spacecraft from LEO to GEO orbits. Even higher powered systems have been proposed to planetary missions such as those shown in Figure 13.54- 55 Having a nuclear vehicle in orbit also enables the beaming of power to air or ground vehicles from orbit or the use of laser and microwave weapons. Figure 13. Nuclear Electric Propulsion (NEP) Concept Vehicles. Left is the 100-kWe Jupiter Icy M oons (JIM O ) planetary spacecraft. Right is a 100-M We-class piloted M ars vehicle. 24 UNCLASSIFIED/Fnn nFFTrinb USE OfiLT UNCLASSIFIED//FOR OFFICIAL UOE ONLY Chapter 4: Summary and Predictions T he principal reason for attractiveness of plasma/M HD propulsion concepts is that they could in principle reach beyond the limits of conventional propulsion, power, and aerodynamic technologies, It is thus near or beyond those limits that novel plasma technologies will likely find their application. T he principal difficulties or flaws associated with the plasma and M HD technologies are as follows: • Weight and complexity, especially if there is a need for a strong (>1 T esla) magnetic field in large volumes, and if electron beams are needed for ionization. • For M HD accelerator/thruster, power requirements could be overwhelming. • M HD operation is accompanied not only by the work of ampere forces, but also by irreversibilities and entropy generation due to Joule dissipation. T his reduces thrust and Isp of propulsion systems. • In the absence of thermal ionization (i.e., at M ach< 12), complexity and power budget associated with nonequilibrium ionization all but make M HD propulsion systems impossible. In contrast, applications to reentry, global-strike hypersonic gliders, and aeroassisted orbital maneuvering look very promising in the near future. T he "free" thermal ionization enables M HD devices with very modest B field and the ability of plasma/M HD system to provide L/D far beyond that possible conventionally; together, it makes these applications both feasible and desirable for national defense. However, these types of applications are also likely to attract attention of other nations, including (but not limited to) Russia, China, and Japan, that have proven knowledge and experience required to accomplish such missions and technologies. T hese nations have the capability to develop such novel technologies within several years and deploying those technologies perhaps within 10 years. T he fortunes of M HD propulsion could increase dramatically if high-speed (hypersonic) vehicles begin to carry powerful onboard electricity sources, such as nuclear (fission or fusion) reactors. Since deployment of onboard nuclear power is mostly a political rather than a technological issue, it is difficult to predict if this going to occur and, if yes, when. For spacecraft, the current trend of replacing chemical rockets with electric propulsion systems will continue and probably will become the standard. Electric systems can provide a much wider range of operation (e.g., low-thrust fine positioning/pointing, more frequent or nontraditional maneuvers, and longer times on station) than chemical systems can. T he trend to larger spacecraft power levels will further accelerate this trend. Ultimately, the high power levels required (hundreds of kW to multi-megawatts) for certain missions will lead to revisiting the use of nuclear fission reactors in space. 25 UNCLASSIFIED/rOR OFFICIAL UOE ONLY U N CLASSI FI E D/ /£OB JQH16iAH^eftEY Chapter 5: Endnotes 1 F. W ilson, “Recent Advances in Satellite Propulsion and Associated Benefits/’ AIAA-20 0 6-530 6, 24th International C ommunications Satellite Systems C onference, San Diego, C A, June 20 0 6. 2 E. C houeiri, “A C ritical History of Electric Propulsion: The First 50 Years,1' Journal of Propulsion and Power, V ol. 20 , N o. 2, M arch-April 20 0 4. ' Journal of Propulsion and Power, Special Section “W eakly Ionized Plasmas for Propulsion Applications,” V ol, 24, N os. 5-6, September-O ctober and N ovember-December 20 0 8. 4 S.O . M acheret, M .N . Shneider, and R.B. M iles, “M odeling of Discharges Generated by Electron Beams in Dense Gases: Fountain and Thunderstorm Regime,” Physics of Plasmas, 20 0 1, V ol. 8, N o. 5, pp. 1518­ 1528. 5 S.O . M acheret, M .N . Shneider, R.B. M iles, and R.J. Lipinski, “Electron Beam Generated Plasmas in Hypersonic M agnetohydrodynamic C hannels,” AIAA Journal, 20 0 1, V ol. 39 , N o. 6, pp. 1127-1136. 6 S.O . M acheret, M .N . Shneider, and R.B. M iles, “M odeling of Air Plasma Generation by Repetitive High- V oltage N anosecond Pulses,” IEEE Transactions on Plasma Science, V ol, 30 , N o. 3, June 20 0 2, pp. 130 1 - 1314. 7 S.O . M acheret, M .N . Shneider, and R.C . M urray, “Ionization in Strong Electric Fields and Dynamics of N anosecond-Pulse Plasmas,” Physics of Plasmas, V ol. 13, 20 0 6, 0 23 50 2. 8 S.O . M acheret, M .N . Shneider, R.B. M iles, and R.J. Lipinski, “Electron Beam Generated Plasmas in Hypersonic M agnetohydrodynamic C hannels,” AIAA Journal, 20 0 1, V ol. 39 , N o. 6, pp. 1127-1136. 9 S.O . M acheret, M .N . Shneider, and R.B. M iles, “M agnetohydrodynamic and Electro hydro dynamic C ontrol of Hypersonic Flows of W eakly Ionized Plasmas,” AIAA Journal, V ol. 42, N o. 7, July 20 0 4, pp. 1378-1387. 10 Journal of Propulsion and Power, Special Section “W eakly Ionized Plasmas for Propulsion Applications,” V ol. 24, N os. 5-6, September-O ctober and N ovember December 20 0 8. 11 S.O . M acheret, M .N . Shneider, and R.B. M iles, “M agnetohydrodynamic and Electro hydro dynamic C ontrol of Hypersonic Flows of W eakly Ionized Plasmas,” AIAA Journal, V ol. 42, N o. 7, July 20 0 4, pp. 1378-1387. 12 E. C houeiri, “A C ritical History of Electric Propulsion: The First 50 Years,” Journal of Propulsion and Power, V ol. 20 , N o. 2, M arch-April 20 0 4. 13 R. Frisbee, editor, “Advanced Space Propulsion C oncepts,” Jet Propulsion Laboratory internal document, January 20 0 2. (This document was accessible via internet until 20 0 3 but has since been removed.) 14 R. Frisbee, editor, “Advanced Space Propulsion C oncepts,” Jet Propulsion Laboratory internal document, January 20 0 2. (This document was accessible via internet until 20 0 3 but has since been removed.) 15 Thrusters, University of M ichigan Plasmadynamics & Electric Propulsion Laboratory, .http://aerospace.engin.umich.edu/spacelab/thrusters/thrusters.html 16 R. Frisbee, editor, “Advanced Space Propulsion C oncepts ” Jet Propulsion Laboratory internal document, January 20 0 2. (This document was accessible via internet until 20 0 3 but has since been removed.) 17 The Lithium Lorentz Force Accelerator for High Power Space Propulsion Project, Electric Propulsion and Plasma Dynamics Lab, Princeton University, .http://alfven.princeton.edu/projects/LiLFA.htm 18 S.O . M acheret, M .N . Shneider, and R.B. M iles, “M agnetohydrodynamic C ontrol of Hypersonic Flow and Scramjet Inlets Using Electron Beam Ionization,” AIAA Journal, V ol. 40 , N o. 1,20 0 2, pp. 74-81. 19 S.O , M acheret, M .N . Shneider, and R.B. M iles, “M HD Power Extraction from C old Air Flow with External Ionizers,” Journal of Propulsion and Power, V ol. 18, N o. 2, 20 0 2, pp. 424-431. Hypersonic 20 M .N . Shneider, S.O . M acheret. and R.B. M iles, “Analysis of M agnetohydrodynamic C ontrol of Scramjet Inlets,” AIAA Journal, V ol. 42, N o. 11, N ovember 20 0 4, pp. 23 0 3-2310 . 21 S.O . M acheret, M .N . Shneider, and R.B. M iles, “O ptimum Performance of Electron Beam Driven M HD Generators for Scramjet Inlet C ontrol.” AIAA Journal. V ol. 45, N o. 9 , 20 0 7, pp. 215 7-2163. 22 B. Parent, S. M acheret, M . Shneider, and N . Harada, “N umerical Study of an Electron-Beam-C onfined Faraday Accelerator,” Journal of Propulsion and Power, V ol. 23, N o. 5, 20 0 7, pp. 10 23-10 32. 23 Kuranov and A. Korabelnikov, “Atmospheric C ruise Flight C hallenges for Hypersonic V ehicles Under the Ajax C oncept,” Journal of Propulsion and Power, V ol. 24, N o. 6, N ovember-December 20 0 8, pp.1229 ­ 1247.’ 26 UNCLASSIFIED/ren OFFICIAL USE ONCT UNCLASSIFIED//FWt OFFICIAL USE ONLY 24 Kuranov and A. Korabelnikov, “Atmospheric C ruise Flight C hallenges for Hypersonic V ehicles Under the Ajax C oncept/' Journal of Propulsion and Power, V ol. 24, N o. 6, N ovember-December 20 0 8, pp.1229 ­ 1247. 25 D. Riggins, “Analysis of the M agnethydrodynamic Energy Bypass Engine for High-Speed Airbreathing Propulsion,” Journal of Propulsion and Power, V ol. 20 , N o. 5, 20 0 4, pp.779 -79 2. 26 D. Riggins, “Analysis of the M agnethydrodynamic Energy Bypass Engine for High-Speed Airbreathing Propulsion,” Journal of Propulsion and Power, V ol. 20 , N o. 5, 20 0 4, pp.779 -792. 27 Kuranov and A. Korabelnikov, “Atmospheric C ruise Flight C hallenges for Hypersonic V ehicles Under the Ajax C oncept,” Journal of Propulsion and Power, V ol. 24, N o. 6, N ovember-December 20 0 8, pp.1229 ­ 1247. 28 Kuranov and A. Korabelnikov, “Atmospheric C ruise Flight C hallenges for Hypersonic V ehicles Under the Ajax C oncept,” Journal of Propulsion and Power, V ol. 24, N o. 6, N ovember-December 20 0 8, pp.1229 ­ 1247. 29 C . Park, U.B. M ehta, and D.W . Bogdanoff, “M agnetohydrodynamics Energy Bypass Scramjet Performance with Real Gas Effects,” Journal of Propulsion and Power, V ol. 17, N o. 5, 20 0 1, pp.10 49 ­ 10 57. 30 C . Park, D.W . Bogdanoff, and U.B. M ehta, “Theoretical Performance of a M agnetohydrodynamic - Bypass Scramjet Engine with N onequilibrium Ionization,” Journal of Propulsion and Power, V ol. 19 , N o. 4, 20 0 3, pp. 5 29 -537. 31 M .N . Shneider and S.O . M acheret, "M odeling of Plasma V irtual Shape C ontrol of Ram/Scramjet Inlet and Isolator," Journal of Propulsion and Power, V ol. 22, N o. 2, 20 0 6, pp. 447-454. 32 M .N . Shneider, S.O . M acheret, R.B. M iles, and D.M . V an W ie, “M HD Power Generation in Scramjet Engines in C onjunction W ith Inlet C ontrol,” AIAA 20 0 4-119 7, 42nd Aerospace Sciences M eeting and Exhibit, Reno, N V , January 5-8, 20 0 4. 33 S.O . M acheret, M .N . Shneider, and R.B. M iles, “Scramjet Inlet C ontrol by O ff-Body Energy Addition: a V irtual C owl,” AIAA Journal, V ol. 42, N o, 11, N ovember 20 0 4, pp. 229 4-230 2. 34 M .N . Shneider, S.O . M acheret, S.H, Zaidi, 1. Girgis, and R.B. M iles, “V irtual Shapes in Supersonic Flow C ontrol with Energy Addition,” Journal of Propulsion and Power, V ol. 25, N o.5, 20 0 8, pp. 9 0 0 -915. 35 S.O . M acheret, M .N . Shneider, and R.B. M iles, “Scramjet Inlet C ontrol by O ff-Body Energy Addition: a V irtual C owl,” AIAA Journal, V ol. 42, N o. 11, N ovember 20 0 4, pp. 229 4-230 2. 36 M .N . Shneider, S.O . M acheret, S.H. Zaidi, I. Girgis, and R.B. 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Shneider, “Three-Dimensional Simulation of the Electric Field and M agnetohydrodynamic Power Generation during Reentry,” AIAA Journal, V ol. 47, N o. 6, 20 0 9 , pp. 1327- 1336. 41 S.O . M acheret, M .N . Shneider, and G.V . C andler, “M odeling of M UD Power Generation O n Board Reentry V ehicles,” AIAA 2 0 0 4-10 24, 42nd Aerospace Sciences M eeting and Exhibit, Reno, N V , January 5-8, 20 0 4. 42 LG. Girgis, M .N . Shneider, S.O . M acheret, G.L. Brown, and R.B. M iles, “C reation of Steering M oments in Supersonic Flow by O ff-Axis Plasma Heat Addition,” Journal of Spacecraft and Rockets, V ol. 43, N o. 3, 20 0 6, pp. 60 7-613. 43 R. Litchford et al., “M agnetohydrodynamic Augmented Propulsion Experiment: I. Performance Analysis and Design,” AIAA 20 0 2-2184, 33rd Plasmadynamics and Laser C onference, M aui, HI, M ay 20 0 2. 44 R. Litchford and J. 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Brophy, “M HD Propulsion Study,” Plasma/Electromagnetic Advanced Propulsion W orkshop, University of Tennessee Space Institute, December 10 -11, 19 9 7. 49 J. Lineberry and J. C hapman, “M HD Augmentation of Rocket Engines for Space Propulsion,” AIAA- 20 0 0 -30 56, 35th Intersociety Energy C onversion Engineering C onference, Las V egas, N V , July 20 0 0 . 50 F. W ilson, “Recent Advances in Satellite Propulsion and Associated Benefits,” AIAA-20 0 6-530 6, 24th International C ommunications Satellite Systems C onference, San Diego, C A, June 20 0 6. 51 V . Hruby et al., “Air Breathing Electrically Powered Hall Effect Thruster,” United States Patent number US 6,834,49 2 B2, December 28, 20 0 4. 52 G. Bennett, “Space N uclear Power: O pening the Final Frontier,” AIAA-20 0 6-419 1,4th International Energy C onversion C onference and Exhibit (IEC EC ), San Diego, C A, June 20 0 6. 53 R. Frisbee, editor, “Advanced Space Propulsion C oncepts,” Jet Propulsion Laboratory internal document, January 20 0 2. (This document was accessible via internet until 20 0 3 but has since been removed.) 54 Jupiter Icy M oons O rbiter (JIM O ) mission web page, Jet Propulsion Laboratory, /http ://w w w2 .j pl .nasa.gov/j imo 55 R. Frisbee, editor, “Advanced Space Propulsion C oncepts,” Jet Propulsion Laboratory internal document, January 20 0 2. (This document was accessible via internet until 20 0 3 but has since been removed.) 28 U NCLASSI FI E D/TOR OFFICIAL UDE ONLT