Direct Fusion Drive For Advanced Space Missions

Transcription

1 Direct Fusion Drive For Advanced Space Missions Mr. Michael Paluszek, Ms. Stephanie Thomas, Mr. Yosef Razin Princeton Satellite Systems, Inc. Prof. Samuel Cohen The Princeton Plasma Physics Laboratory Dr. David Farley Quantal Technologies, LLC Princeton Satellite Systems, Inc. 6 Market Street Street, Suite 926 Plainsboro, NJ Phone: Fax: map@psatellite.com Web:

2 Executive Summary Princeton Satellite Systems and the Princeton Plasma Physics Laboratory (a Federally-Funded Research and Development Center) are collaborating on a breakthrough concept, the Direct Fusion Drive (DFD), a D 3 He fueled fusionpowered rocket engine. The design is based on the Princeton Field-Reversed Configuration Reactor (PFRC) concept for a compact, clean, steady-state fusion reactor for power levels between 1 and 10 MW. Larger power levels can be accommodated using multiple engines. This technology enables missions such as human Mars orbiters, nuclear fusion powered space stations and high power robotic missions to the outer planets. This paper reviews the theory behind the design of compact and clean (aneutronic) fusion-powered rocket engines. The reactor employs a field-reversed magnetic field configuration (FRC) for plasma confinement, a novel radio-frequency (RF) method for plasma heating, and a unique method for control of propellant flow, producing a thrust of 30 N at a specific impulse of s. The FRC has higher β (plasma pressure/magnetic energy density) than other magnetic plasma-confinement devices and a relatively simple linear solenoid magnet layout that permits control of propellant flow parameters. Higher β reduces the mass of the superconducting solenoidal coils needed for achieving the higher plasma temperatures required for aneutronic fusion reactions. Waste heat generated from the plasma s bremsstrahlung and synchrotron radiation can be recycled to power the RF heating and other on-board systems. The unique characteristics of this reactor with non-equilibrium plasma conditions will produce five-fold lower levels of neutrons than from equilibrium conditions. Additionally, the reactor design permits a novel method for extracting electric energy directly from the fusing plasma using the RF system. To produce fusion power with few neutrons, the reactor employs the D 3 He reaction, which requires plasma temperatures about five times greater than the deuterium-tritium (D T) reaction. A profound benefit results: the D 3 He reaction can produce 20-times fewer and 5-times lower-energy neutrons than the D T reaction per unit of power produced. Combined with the small size of the reactor, and its high surface-to-volume ratio, substantially less neutron-shielding mass is needed compared to mainstream D T burning tokamak reactor designs; longer component lifetimes result. A sequence of two FRC test devices, followed by a prototype rocket engine, can be designed, built, and tested in fouryear steps compared to the 30-year timeframe for design and construction of a single large D T burning tokamak, like ITER, still two steps away from a practical reactor. The PFRC approach would permit more rapid and less costly testing, evaluation, and improvement. The twelve-year research program that would result in a D 3 He burning FRC prototype engine that demonstrated the critical spacecraft-propulsion technologies would cost about $76M. 1

3 Contents 1 The Princeton Field-Reversed Configuration Reactor Introduction The Field Reversed Configuration The RMF o Heating System System Design Engine Overview Shielding RMF Heating Magnetic Nozzle and Thrust Augmentation Systems Thermal Power Conversion Radiators Startup The Helium-3 Supply 8 4 On-Going Research PFRC Magnetic Nozzle Trajectory Design and Balance of Plant Related Patent Applications Application to Space Propulsion Overview Robotic Mission to Sun-Earth L Human Missions to Mars Program Plan Integrated Development Plan References 13 2

4 1 THE PRINCETON FIELD-REVERSED CONFIGURATION REACTOR 1 The Princeton Field-Reversed Configuration Reactor 1.1 Introduction The Princeton Field-Reversed Configuration Reactor (PFRC-R) would be a 2 m diameter, 10 m long, steady-state plasma device heated by a novel radio-frequency (RF) plasma-heating system, enabling the achievement of sufficiently high plasma temperatures for D 3 He fusion reactions. An FRC employs a linear solenoidal magnetic-coil array for plasma confinement and operates at higher plasma pressures, hence higher fusion power density for a given magnetic field strength than other magnetic-confinement plasma devices. A linear solenoid is well-suited for producing a collimated directed exhaust stream that may be used for propulsion. A rocket engine based on the PFRC is designed to operate with a D 3 He fuel mixture though, for decade-long missions, it may be operated with a tritium-suppressed D D fuel cycle. Both fuels produce much lower levels of neutrons than deuterium-tritium, reducing shielding mass as well as waste energy unavailable for propulsion. In the PFRC-R, waste heat generated from bremsstrahlung and synchrotron radiation will be recycled through the RF system to maintain the fusion temperature. The features of this design are: 1. Odd-parity rotating magnetic field (RMF) heating for high stability and efficient heating 2. Non-equilibrium operation for reduced neutron production 3. Can operate with D 3 He or tritium-suppressed D D 4. Electric power extraction through the RMF coils 5. Combined thrust and electric power generation 6. Variable specific impulse and thrust through deuterium augmentation in the scrape-off layer 7. High temperature superconductors for plasma confinement and the magnetic nozzle that result in drastically reduced cooling requirements 8. Engines are from 1-10 MW in size which is an ideal range for space power and propulsion 9. Multiple engines can be combined to produce higher levels of thrust and power The following sections provide more detail on the PFRC. 1.2 The Field Reversed Configuration An important figure-of-merit for fusion reactors is β, the ratio of the plasma pressure to the magnetic energy density. Of all candidate magnetic-confinement fusion reactors, FRCs have the highest β, see [1] for a review of early FRC research. Accordingly, FRC magnets would be less massive than those for a tokamak of comparable power. High β is essential for burning aneutronic fuels, such as D 3 He, because they require higher ion energies to achieve the same fusion reactivity as D T. FRC plasma-confinement devices have at least two other attractive features, notably, a linear magnet geometry [1] and a natural divertor. This structure provides an ideal attachment point for a magnetic nozzle, allowing for the control of the plasma exhaust and its plume angle, for use as a propulsive or power-producing device. The FRC is unique among quasi-toroidal, closed-field-line magnetic confinement devices in that it is simply connected. FRCs also have zero toroidal magnetic field, no internal conductors, and a line of zero magnetic field strength within the plasma encircling its major axis, termed the O-point null line. This O-point null line proves essential for our proposed method of RF plasma heating. Figure 1-1 on the following page shows the FRC s magnetic-field structure and the linear coil array. A separatrix divides a closed-field region (CFR) from the open-field region (OFR). Fieldshaping coils that are magnetic flux conservers surround the plasma. The O-point null line, not shown, is a co-axial ring located on the magnetic axis, the center of the nearly elliptical CFR. The reactor design we propose differs from that of Cheung [2] in size, heating method, and fuel. Cheung et al. selected p 11 B, which requires five-times higher ion energies and produces far less fusion power per reaction. Cheung et al. selected neutral beams for heating, requiring a plasma volume that is one hundred times larger, is therefore more costly 3

5 1.3 The RMF o Heating System 1 THE PRINCETON FIELD-REVERSED CONFIGURATION REACTOR Figure 1-1: DFD Core showing the details of the PFRC reactor and the linear path of the propellant flow. Closed Field Region Box Coil Field Shaping Coils Separatrix Nozzle Coil Gas Box Exhaust Plume Propellant SOL Heating Section D-He3 Fusion Potential Drop Ion Acceleration Region and less stable. The heating technique selected in Miller [3] is called an even-parity rotating magnetic fields (RMF e ) [4], a method that has shown poor energy confinement, and requires larger FRCs. FRCs where the plasma radius is more than 10 times their ion Larmor radius are prone to magnetohydrodynamic (MHD) instabilities. To achieve better energy confinement, we instead invented odd-parity RMF, RMF o, allowing for smaller, more stable reactors. Many physics challenges remain before the RMF o heated FRC can be developed into a practical reactor. The predictions of excellent energy confinement and stability and of efficient electron and ion heating to fusion-relevant temperatures, must be validated. Substantial progress has occurred in the first three areas. In 2010 and 2012, TriAlpha Energy Corp reported near-classical energy confinement time in their FRC [5, 6]. (Classical confinement time occurs for Coulomb-collision-driven diffusion only. The confinement time of real plasma is often far less than the classical limit [7].) Our reactor needs energy confinement only 1/5 as large as the classical. In 2007, an RMF o heated FRC [8] achieved stable plasma durations 3,000 times longer than predicted by MHD theory [9]; by 2012 that record was extended to over 10 5 times longer. Finally, theoretical studies [10, 11, 12] indicate that RMF o will be able to heat plasma electrons and ions to fusion relevant temperatures. These are promising starts, but much research is needed at higher plasma temperatures and densities, and with burning, i.e. fusing, plasmas. Based on 1/5-classical-confinement, a plasma radius of 25 cm is adequate for confining the high energy plasma needed to produce 1 MW of fusion power. This radius matches criteria set by the RMF o heating method. Modest changes in parameters could increase the fusion power up to 10 MW. 1.3 The RMF o Heating System For an FRC reactor to burn a D 3 He fuel mixture, the plasma ions must be heated to over 50 kev. If energetic neutralbeam injection were used for achieving these temperatures, the plasma would have to be over 4 m in diameter in order to stop the injected neutral beam. Such a large reactor would produce proportionally large amounts of power, near 1 GW. In RF heating methods, on the other hand, power can be absorbed over shorter distances. Using RMF o allows the diameter of the plasma to be reduced to 0.5 m and produce 1 MW. In RMF o heating, the maximum ion energy is proportional to the RMF o frequency ω RMF. Due to a constraint set by the RMF-generated current and the FRC s magnetic field strength, ω RMF decreases as the product of plasma density times the square of the plasma radius. Thus, too large or dense an FRC is not well heated. An optimum FRC for RMF o heating of ions to 100 kev and above has a radius in the range cm. This, naturally, places a lower limit on the confinement time required, no worse than a fifth of the classical value, as noted earlier. RMF o creates a time-varying azimuthal electric field near the O-point null line [10]. This periodically accelerates ions 4

6 2 SYSTEM DESIGN into betatron orbits and then decelerates them back into cyclotron orbits. Choosing the RMF o ω RMF and amplitude properly allows ions to be pumped up, repeatedly, to energies near the peak of the D 3 He fusion cross-section and then returned to the bulk temperature. This is a conservative process that satisfies the recirculating energy criterion derived by Rider [13] to sustain, against collisions, a non-maxwellian distribution that increases the fusion rate and decreases neutron production. In a D 3 He plasma, the trajectories of RMF o accelerated ions are predicted to form two betatron-orbit streams close to the FRC s O-point null line, a D stream and a 3 He stream. The formation of these overlapping streams reduces f p, the fraction of fusion energy in neutrons. The deuterium stream ions have half the peak energy of the 3 He ions, causing non-zero relative velocity between them. The energy dependent fusion rates can be used to show the basic effect of the higher energy of the 3 He stream. If the bulk plasma has an average energy of 70 kev and the RMF pumps the 3 He up by 100 kev it will pump the deuterium up by only 50 kev. Thus, several effects centrally peaked betatron orbits, low transverse beam temperature, reduced D/ 3 He ratio, and higher 3 He energy combine to decrease f p from 33% to below 0.5% for an RMF o heated D 3 He -fueled FRC. Add to this that D 3 He produced neutrons have only around 1/6 of the energy of those produced by D T and the larger surface-to-volume ratio ( 1/radius) for a small (25 cm) FRC compared to a large (10 m) tokamak and an additional 200-fold reduction of neutron power load per unit area on the wall is obtained. Overall, the shielding requirements for this type of small, clean reactor are far less than for a D T -fueled larger fusion engine. The basic design for a PFRC reactor s neutron shield is discussed in Section 2.2. The RMF o method, in addition to reducing neutron production, also offers the possibility of a novel, direct energy extraction method from the fusion products. The same rotating azimuthal electric field that heats the ions can be used to extract energy from the 3.6 MeV alphas and 14.7 MeV protons produced by D 3 He fusion. 2 System Design 2.1 Engine Overview The engine overview is shown in Figure 4-1 on page 9 including the balance of plant components needed for a space mission. The major subsystems are the 1. PFRC reactor 2. Shielding 3. RMF heating 4. Magnetic nozzle 5. Superconducting coils and refrigeration 6. Thermal power conversion 7. Radiators 8. Startup power The subsystems are discussed below. 2.2 Shielding Even though D 3 He is an aneutronic fuel, one of its D-D side reactions does produce 2.45 MeV neutrons. As neutron absorption cross-sections are many orders of magnitude smaller than their scattering cross-sections, the neutrons must be slowed significantly for attenuation to occur. Findings from Project Prometheus [14] and [15] showed that the best materials for neutron attenuation and absorption were low-z (low atomic number) materials such as H (possibly in water), Li, Be, and B, particularly 10 B. Liquid water is an excellent material for neutron shielding, but difficult 5

7 2.3 RMF Heating 2 SYSTEM DESIGN Figure 2-1: DFD System design for a space mission. Spacecraft Power Bus RF RF Generator Space D 0 Injector Magnetic Nozzle Coil Refrigerator Radiator 3 He Injector Plasma HTS Coil Blanket RF X-ray n Helium Coolant Generator Radiator 3 He Shielding Heat Engine Power Recycling Restart D 2 O D APU O 2 Electrolysis to handle in low pressure and low temperature environments. For space applications, solid materials are preferable. Solid low-z material may also serve as liners for the reactor chamber as low-z elements cause less problematic plasma contamination. These liners must be able to handle the extreme environment of space and also the internal thermal, electrical, and structural conditions in the reactor. 10 B strongly attenuates neutron energy and, unlike LiH and Be, is also a neutron absorber and therefore reduces the total flux behind the shield. As 10 B 4 C it is a very strong, stable solid over a very large range of temperatures. 10 B 4 C is readily available in nature, safe to manufacture, inexpensive, and it does not exhibit any obvious structural deformities when irradiated. Since it is transparent to x-rays, bremsstrahlung radiation can pass through it to a thermal power conversion system. 10 B 4 C is our first choice for shielding and 10 BN our second. X-ray shielding is also required. Many materials absorb X-rays. Tungsten is one of the best attenuators on a per mass basis [14] and meets many of the material requirements for shielding including availability, manufacturability, and high temperature capabilities. Thin tungsten sheet would be placed in channel inside the 10 B 4 C shielding. helium coolant would extract from the heat deposited by absorbed X-rays in the tungsten. 2.3 RMF Heating The RF subsystem consists of the RF generator, coaxial cables, directional couplers, tank circuits, antennas and RF controls and diagnostics. The RMF o heating system generates RF power somewhat below the 3 He cyclotron frequency that falls within the medium- to high-frequency radio bands, about MHz. The RF system must produce enough power to start and sustain the fusion reactions. When the fusion reactor is operating, smaller amounts of RF power are needed because fusion products provide heating and current drive. The PFRC-2 discussed in Section 4.1 has a 200 kw RF system; we estimate a peak RF power of 1 MW system will be needed for a 5MW PFRC-R. 6

8 2.4 Magnetic Nozzle and Thrust Augmentation Systems 2 SYSTEM DESIGN 2.4 Magnetic Nozzle and Thrust Augmentation Systems The magnetic nozzle and thrust augmentation system allow direct control of the thrust level and exhaust velocity. The magnetic nozzle [16, 17, 18] redirects the flow from the FRC to free space. The nozzle consists of a throat coil and two or more additional nozzle coils to allow expansion and acceleration of the flow. Methods to control the plume angle are under investigation. All of the coils are superconducting. Thrust is increased by the addition of propellant, increasing mass flow and thrust [19]. In our design deuterium gas is introduced at the top the far left side of the engine; see Figure 1-1 on page 4. The gas would be ionized quickly, forming a plasma similar to what is called, in tokamak terminology, a high recycling divertor. The cold plasma will flow to the right, with electrons picking up energy from the fusion products and synchrotron radiation as it passes over the ellipsoid. The shell plasma thus formed and heated will be quite similar to that already produced in gas dynamic traps [20], having a density near cm 3 and an electron temperature of ev. The ions will be cooler. As the plasma leaves the reactor, acceleration of the ions to an energy near 4T e will occur in a double layer [21] near the nozzle throat. Within a few centimeters beyond the nozzle throat, the propellant will have acquired the desired kev energy. 2.5 Thermal Power Conversion As noted above it is possible to extract energy from the plasma using the RF system. Thermal power from synchrotron radiation that heats the structural shell and bremsstrahlung radiation and heat from the neutrons hitting the shielding would be converted to electrical power using a heat engine. Stirling [22], Brayton cycles [23] and direct conversion methods [24] are possibilities. Modern gas turbines have specific powers of 10 kw/kg. This number should be achievable in this system. The bremsstrahlung can be efficiently removed by flowing helium gas through B 4 C heat exchanger tubes containing corrugated sheets of refractory medium-z material [25]. B 4 C, beyond being a good neutron shield, is an insulator and nearly transparent to 100 kev X-rays; high-z materials such as tungsten absorb the X-rays within a few millimeters. The bremsstrahlung heats the high-z materials to about 2,000 K. The helium coolant extracts this energy and is used to drive a high-efficiency thermal cycle. With a thermal rejection temperature to space of 650 K, this results in a Carnot efficiency of 65% [26]. 2.6 Radiators Carbon-carbon composite radiators are under development for dissipating waste heat from spacecraft [27]. The radiator is bonded to metal heat pipes and operates in conjunction with them at a temperature between 500 and 1,000 K. The areal mass of this type of radiator is calculated to be 2.2 kg/m 2, almost a factor of 5 better than current radiators. 2.7 Startup The DFD needs the ability to restart in space. Other proposed large fusion engines [28] have included fission reactors for startup. For DFD an auxiliary power unit provides startup power. The combustion of deuterium and oxygen produces heat that is used by a heat engine to produce electricity. This can be supplemented by electricity from other operating engines. The exhaust from the combustion is condensed and electrolyzed to produce hydrogen and oxygen once the engine is in operation. This provides a constant source of fuel for future startups. The engine is started up at partial power in electricity generation mode and this power replaces the power from the combustion as the fusion power increases. The combustor uses the same heat engine as the fusion engine uses for power generation. 7

9 4 ON-GOING RESEARCH 3 The Helium-3 Supply 3 He is scarce on Earth s surface and demand is high. In 2010, demand for 3 He was projected to be 76,000 liters (at Standard Temperature and Pressure about 10 kg), per year, but the United States only produces 8,000 liters of 3 He a year. Moreover, last year the reported U.S. stockpile of 3 He was at less than 48,000 liters [29]. 3 He currently costs $100 USD/liter. It is estimated that 26,000 liters of 3 He could be produced per year as a byproduct of natural gas and helium production [29]. This would cost between $34 and $300 USD/liter. A plant to extract 3 He would cost several million dollars. The cost is based on extracting it from helium that is already produced. If 3 He were the sole product of the plant, the cost would be $12,000 USD/liter. The atmosphere contains approximately 280 billion liters of 3 He but it is uneconomical to extract it from the air. Increased production of tritium, which decays to 3 He, in fission power plants is another option. The readily accessible terrestrial 3 He would sustain 100 MW of continuous power production for about 100 years. Based on this supply only, 3 He is feasible for space propulsion and selected military applications. Helium-catalyzed D-D fusion would avoid the 3 He shortage problem but would require additional neutron shielding, an approach that merits further study. 4 On-Going Research 4.1 PFRC-2 The PFRC-2 project is in its second year of a three-year grant from DOE, a continuation of PFRC grants from DOE that date back to The PFRC research program explores effects of odd-parity rotating magnetic fields RMF o on the basic plasma science topics of FRC formation, current drive, electron and ion heating, plasma confinement, and plasma stability. The PFRC s goal is to achieve, within two years, stable, kev-temperature plasmas of 0.1 s duration. (Already hundreds of stable lower temperature plasmas of 0.18 s duration have been created.) Program success, aimed at the eventual development of advanced-fuel operational scenarios, will have a strong impact on the design of a practical fusion reactor. The research is being performed in collaboration with scientists at the Los Alamos National Laboratory (LANL), Lawrence Livermore National Lab (LLNL), University of Washington, ETH Zurich, and Voss Scientific, LLC. It provides excellent education and training opportunities to graduate and undergraduate students. 4.2 Magnetic Nozzle The Magnetic Nozzle (MNX) project is in the first year of a three-year grant from DOE and is a continuation of a research effort begun in MNX comprises a set of experiments and theoretical efforts that investigate phenomena predicted to occur in warm plasmas flowing on open magnetic field lines that bound hot plasma in closed field line regions, e.g., FRCs. The primary research areas are plasma detachment in the exhaust plume, field-parallel potential drops near the magnetic nozzle, and non-local cross-field heat transport from the closed-field-line region into the openfield region. The experiments are performed in the MNX device, part of the PFRC-2 facility. The PFRC-2 has recently been upgraded to 10 higher pumping speed in the plasma expansion region and to 100-times higher heating power, to 200 kw. Theoretical efforts bring an extensive and more representative set of physics equations to this research by augmenting fluid (UEDGE) modeling with kinetic (LSP) simulations, to be performed by a multi-institutional team (including LLNL and Voss Scientific) with extensive student involvement. MNX is a member of the class of open-field-line linear plasma devices with warm (T e = ev) plasma flowing out both ends through mirror coils. Intense gas puffing is applied into a coaxial closed box beyond one mirror coil, Figure 1-1 on page 4, while an outflow of plasma into vacuum is permitted beyond the second mirror coil. This configuration is under consideration for advanced spacecraft propulsion schemes, new plasma processing tools, and other applications. MNX will investigate the parameter range of simultaneous plasma detachment at both ends, by plasma cooling and recombination in the gas box and by super-alfvenic flow in the expanding field region, the latter a game changing 8

10 4.3 Trajectory Design and Balance of Plant 4 ON-GOING RESEARCH prediction for advanced rocket engines. Theoretical studies will explore whether non-local cross-field heat transport, particularly by fast ions, will create broad and uniform heating across the open-field-line scrape-off layer (SOL) and whether the SOL thickness and parameters can be controlled, to allow more uniform exhaust of momentum and power. We will explore fast ion slowing down in a regime not previously investigated, where the ion motion is perpendicular to the magnetic field, the ion velocity greater than the electron velocity. Figure 4-1: Experiments in progress showing PFRC-2 test and MNX testing. 4.3 Trajectory Design and Balance of Plant Princeton Satellite Systems IR&D is funding work on trajectory design and research into the balance of plant components for the fusion rocket engine employing the PFRC-R. The trajectory research is developing algorithms and software to take advantage of the variable exhaust velocity and thrust of the engine for both unmanned and manned missions [30, 31, 32, 33]. The balance of plant work includes work on shielding, electric power generation, RF heating systems, radiators, superconducting coils and structural materials. 4.4 Related Patent Applications The following are joint Princeton Satellite Systems and Princeton Plasma Physics Laboratory patent applications related to DFD. M. Buttolph, D. Stotler and S. Cohen, Fueling Method for Small, Steady-State, Aneutronic FRC Fusion Reactors, 61/873,651, filed September 4, M. Paluszek, E. Ham, S. Cohen and Y. Razin, In Space Startup Method for Nuclear Fusion Rocket Engines, 61/868,629, filed August 22, S. Cohen, G. Pajer, M. Paluszek and Y. Razin, Method to Reduce Neutron Production in Small Clean Fusion Reactors, PCT/US13/33767, filed March 25, S. Cohen, G. Pajer, M. Paluszek and Y. Razin, Method to Produce High Specific Impulse and Moderate Thrust From a Fusion-Powered Rocket Engine, PCT/US2013/40520, filed May 10,

11 5 APPLICATION TO SPACE PROPULSION 5 Application to Space Propulsion 5.1 Overview The future of space exploration, from robotic deep-space expeditions to manned interplanetary missions, will require high thrust and high exhaust velocity (u e ) engines. The cost of these missions can be reduced by increasing these parameters, thereby decreasing mission transit time and the mass of the spacecraft. This is especially crucial for deep-space missions, partly because their operational costs can approach $50M USD per year. For manned missions, reducing transit time has the additional benefit of reducing the astronauts exposure to cosmic radiation and low gravity. Furthermore, high-power propulsion is required for manned missions in the case of an aborted operation; this allows the astronauts to quickly return to Earth after an accident or emergency. The exact engine specifications, particularly power and specific impulse (I SP ), will vary for each mission. Many proposed NASA and ESA missions require high-performance propulsion systems, as shown in Table 5-3. All of these goals are exceeded by the power output of a single PFRC-R, which can potentially achieve up to 10 MW of power. Thus, the PFRC-R, theoretically meets the anticipated power requirements for deep-space and interplanetary manned missions. The advantage in specific power of the PFRC-R over other technologies is illustrated in Table 5-2. The low specific powers for nuclear electric make high-power thrusters, like MPD, LiFo or VASIMR, uncompetitive except in the inner solar system where light-weight solar arrays have specific powers above 0.1. PFRC-R could be used to produce electricity for such engines but the overall specific power would be lower than a PFRC-R that produced thrust by itself. Therefore, the PFRC-R promises the highest specific power of the alternatives. Table 5-1: Proposed space missions requiring high power propulsion systems. Mission Power (kw) Power Source Engine Reference JIMO 180 Nuclear Fission Nexis Ion and Hall [33] Thrusters Outer Planets 95 Nuclear Fission Nexis Ion [33] 200 AU 65 Nuclear Fission DS4G [31] 200 AU 160 Solar Panels Ion [31] NEO 2004 MN4 210 Solar Panels Ion [32] NEO Crew 350 Solar Panels Hall [32] Table 5-2: Space Power Plants giving specific powers for several technologies. Type kw/kg Ref FRC 1 [34, 35, 36] Solar Electric 0.2 [31] Nuclear Electric 0.04 [31] Radioisotope [37] Table 5-3: High Power Thrusters DFD produces higher thrust and power. Thruster Power (kw) Thrust (N) Specific Impulse (s) Reference VASIMR [38] NASA-457M to 2740 [39] LiFO [40] DFD to [34] 5.2 Robotic Mission to Sun-Earth L2 An operating point for a nominal 1 MW PFRC-R rocket engine can be selected; a plasma radius of 20 cm is adequate for confining the high energy plasma needed to produce 1.5 MW of fusion power. The plasma temperatures listed are approximations for the full non-maxwellian particle distributions generated and sustained by the RMF o. Figure 4-1 on the preceding page depicts the overall power and propulsion system. The RF system will have to supply about a 10

12 5.3 Human Missions to Mars 5 APPLICATION TO SPACE PROPULSION 1 MW burst of power for forming and heating the FRC to fusion temperatures. After that, the RMF o power can be reduced by a factor of two. The example mission is for a spacecraft to go from low earth orbit to the Sun-Earth L2 point to deploy the James Webb Space Telescope (JWST). The spacecraft would enter a 750,000 km halo orbit about L2 [41]. The spacecraft acts as a ferry and returns to low-earth orbit for restocking (including propellant) and further missions which might include servicing the JWST. The thrust and exhaust velocity can be traded against each other. For any particular total change in velocity, V, there will be an optimal u e that minimizes the mission mass. For this mission the thrust was set to 20 N and the exhaust velocity to 54.6 km/s. An example trajectory is shown in Figure 5-1. This trajectory is simulated in a non-dimensional, circular restricted three-body problem (CRTBP) for the Sun-Earth/Moon system. The spacecraft is initialized in a km Earth orbit with zero inclination. Thrust is applied for approximately 12.7 days in a direction normal to the plane defined by the Earth-relative position vector and the orbit-normal vector (z). This causes the spacecraft to gradually increase the altitude of its Earth orbit as well as its eccentricity, spiraling out until it escapes Earth gravity. This is followed by a coast arc of 16 days and then a 16 hour insertion burn, allowing the spacecraft to coast into a weakly stable halo orbit about L2. This trajectory was simulated with multiple iterations, adjusting the thrust to control mission duration, and changing the fuel mass and associated dry mass until the required V was feasible. Figure 5-1: DFD mission to L2 showing an example trajectory Lunar Orbit Insertion Burn at 28.5 days (for 16 hrs) 37.5 days Y (Thousands km) L Spiral from GPS (12.7 days) Delta V = 3.1 km/s Initial Alt. = km From LEO: Add 3.6 km/s, 15.9 days X (Thousands km) 5.3 Human Missions to Mars Human missions to the planets have been planned since before the engine of the Apollo program in the early 1970 s. At that time the most advanced propulsion option was nuclear fission thermal rockets which heat hydrogen flowing through a reactor core. Several of these engines were tested in the late 1960 s and early 1970 s, before the fission engine programs were cancelled. Recent work reports that the radiation data the Curiosity rover collected on its way to Mars found astronauts traveling to and from Mars would be bombarded with as much radiation as they d get from a full-body CT scan about once a week for a year. [42]. Add to that the harmful effects of muscle atrophy from long-term low-gravity, and mission speed becomes a clear priority to ensure the crew s health. Consequently, chemical or nuclear thermal rocket transfers would not be sufficient for human exploration to Mars and more distant destinations. The DFD powered transfer stage can get astronauts to Mars in months, significantly reducing radiation exposure and atrophy effects. To demonstrate the potential of DFD technology, we developed a concept for a Mars orbital mission that uses NASA s 11

13 5.3 Human Missions to Mars 5 APPLICATION TO SPACE PROPULSION Deep Space Habitat to house the crew [43]. With a variable thrust augmentation system, the DFD allows for power generation up to 10 MW, which is ideal for interplanetary exploration. The Orion spacecraft would launch the astronauts into low Earth orbit where it would dock with the DFD transfer vehicle. The baseline vehicle is shown in Figure 5-2 docked with the Orion. It has five 6 MW engines. This provides redundancy and an abort capability in case of an engine failure. The engines provide both propulsion and electric power during the mission. Figure 5-2: Orion spacecraft docked with the Deep Space Habitat on the DFD transfer vehicle. The round-trip mission to Mars, with a spacecraft powered by DFDs and carrying NASA s Deep Space Habitat, involves two orbit transfers which involve a long burn, a coasting period and another long burn. The spacecraft enters and departs Earth and Mars orbits using the DFD. No aerodynamic braking is required. This double-rendezvous problem typically requires waiting for a full synodic period (i.e. the next time the two planets return to their current alignment), which is 780 days for Earth and Mars. The goal, however, is to make this roundtrip as quickly as possible. Therefore, PSS calculated a new roundtrip trajectory that takes a short-cut of sorts, traveling inside Earth s orbit on the return flight. This enables the overall mission to be shortened to just 310 days, including a 30 day stay at Mars. Figure 5-3: Round trip mission to Mars. takes only 310 days including 30 days in Mars orbit. DeltaV(km/s) = 50, Trip Time(days) = Coast Trajectory Burn Trajectory 1 Mars Orbit 0.5 Y (AU) 0 Earth Orbit Sun X (AU) Princeton Satellite Systems The Mars mission uses the NASA Space Launch System for launch into low Earth orbit. NASA s Space Launch System is an advanced, heavy-lift launch vehicle which will provide an entirely new capability for science and human exploration beyond Earth s orbit. The SLS has two variations, 70 MT and 130 MT. The Mars orbital mission will require one 130 MT SLS launch, as the spacecraft weighs 129MT. A separate launch brings the crew to the spacecraft in an Orion spacecraft. On-orbit testing and checkout is done prior to the arrival of the crew. 12

14 REFERENCES Figure 5-4: SLS 130 MT variant with the Mars spacecraft superimposed. 117 m 6 Program Plan 6.1 Integrated Development Plan PSS and PPPL have developed a complete DFD development program designed to produce an operating fusion reactor within a 12-year time frame. The current work is in the experimental stage and is TRL 3-4. Two additional research devices would be built, PFRC-3 and PFRC-4. PFRC-4 would demonstrate fusion power generation. It is estimated that this would cost $76M. This development plan is much shorter and the cost much less than that for Tokamak fusion. This is due to the simple geometry and the small size of the machines. Furthermore, the experimenters do not have to handle tritium, an expensive process. Figure 6-1: DFD Program Plan resulting in a burning fusion reactor by i-temp is ion temperature and e-temp is electron temperature. Machine Objectives PFRC-1 PFRC-2 PFRC-3 PFRC-4 Electron Heating Ion Heating Heating above 5 kev D-He3 Fusion Goals/ Achievements* 3 ms pulse* 0.15 kg field* e-temp = 0.3 kev* 0.1 s pulse* 1.2 kg field i-temp = 1 kev 10 s pulse 10 kg field i-temp = 5 kev 1000 s pulse 60 kg field i-temp = 50 kev Plasma Radius 4 cm 8 cm 16 cm 25 cm Time Frame Total Cost $2M $6M $20M $50M References [1] Tuszewski, M. A., Field Reversed Configurations, Nuclear Fusion, Vol. 28, No. 11, November 1988, pp [2] Cheung, A., Binderbauer, M., Liu, F., Qerushi, A., Rostoker, N., and Wessel, F. J., Colliding Beam Fusion Reactor Space Propulsion System, Space Technologies and Applications International Forum, No. CP699, [3] Miller, K., Slough, J., and Hoffman, A., An Overview of the Star Thrust Experiment, Space technology and applications international forum, Vol. 420, AIP Conference Proceedings, , pp [4] Blevin, H. and Thonemann, P., Proceeding of the Conference on Plasma Physics and Controlled Nuclear Fusion Research, Nuclear Fusion Supplement, Vol. 1, 1962, pp. 55. [5] Binderbauer, M., Guo, H. Y., M.Tuszewski, Putvinski, S., Sevier, L., and Barnes, D., Dynamic Formation of a Hot Field Reversed Configuration with Improved Confinement by Supersonic Merging of Two Colliding High-Compact Toroids, Physical Review Letters, Vol. 105, 2010, pp

15 REFERENCES REFERENCES [6] Tuszewski, M., Smirnov, A., and Thompson, M., Field Reversed Configuration experiment through edgebiasing and neutral beam injection, Physical Review Letters, Vol. 108, No , June [7] Wesson, J., Tokamaks, Oxford Univ. Press, Oxford, 3rd ed., [8] Cohen, S. A., Berlinger, B., Brunkhorst, C., Brooks, A., Ferarro, N., Lundberg, D., Roach, A., and Glasser, A., Formation of Collisionless High-β Plasmas by Odd-Parity Rotating Magnetic Fields, Physical Review Letters, Vol. 98, 2007, pp [9] Rosenbluth, M. N. and Bussac, M. N., MHD Stability of Spheromak, Nuclear Fusion, Vol. 19, 1978, pp [10] Glasser, A. and Cohen, S. A., Ion and electron acceleration in the Field-reversed configuration with an oddparity rotating magnetic field, Physics of Plasmas, 2002, pp [11] Landsman, A. S., Cohen, S. A., and Glasser, A., Onset and saturation of ion heating by odd-parity rotating magnetic fields in an FRC, Physical Review Letters, Vol. 96, 2006, pp [12] Cohen, S. A., Landsman, A. S., and Glasser, A., Stochastic ion heating in a field-reversed configuration geometry by rotating magnetic fields, Physics of Plasmas, Vol. 14, 2007, pp [13] Rider, T., Fundamental limitations on plasma fusion systems not in thermodynamic equilibrium, Physics of Plasmas, Vol. 4, 1997, pp [14] Lewis, R., Space Shielding Materials for Prometheus Application, Tech. Rep. MDO , Knolls Atomic Power Laboratory, January [15] Bushman, A., Carpenter, D. M., Ellis, T. S., Gallagher, S. P., Hershcovitch, M. D., Hine, M. C., Johnson, E. D., Kane, S. C., Presley, M. R., Roach, A. H., Shaikh, S., Short, M. P., and Stawicki, M. A., The Martian Surface Reactor: An Advanced Nuclear Power Station for Manned Extraterrestrial Exploration, Tech. Rep. MIT-NSA- TR-003, MIT Nuclear Space Applicaion Program, December [16] Arefiev, A. V. and Breizman, B. N., MHD scenario of plasma detachment in a magnetic nozzle, Tech. rep., Institute for Fusion Studies, The University of Texas, July [17] Tarditi, A. G. and Steinhauer, L. C., Plasma Flow Control in a Magnetic Nozzle for Electric Propulsion and Fusion Scrape-Off Layer Applications, 2011 International Sherwood Fusion Theory Conference, No. 2C31, November [18] Cohen, S., Sun, X., Ferraro, N., Scime, E., Miah, M., Stange, S., Siefert, N., and Boivin, R., On collisionless ion and electron populations in the magnetic nozzle experiment (MNX), IEEE Transactions on Plasma Science, Vol. 34, No. 3, june 2006, pp [19] Santarius, J., Lunar He-3, fusion propulsion, and space development, Lunar Bases and Space Activities of the 21st Century, edited by W. Mendell, J. Alred, L. Bell, M. Cintala, T. Crabb, R. Durrett, B. Finney, H. Franklin, J. French, and J. Greenberg, Sept. 1992, pp [20] Anikeev, A., Bagryansky, P., Ivanov, A., Karpushov, A., Noack, K., and Strogalova, S., Upgrade of The Gas Dynamic Trap: Physical Concepts and Numerical Models, 28th EPS Conference on Controlled Fusion and Plasma Physics, Funchal, Vol. ECA, 25A, 2001, p [21] Cohen, S., Siefert, N., Stange, S., Boivin, R., Scime, E., and Levinton, F., Ion acceleration in plasmas emerging from a helicon-heated magnetic-mirror device, Physics of Plasmas, Vol. 10, 2003, pp [22] Dhar, M., Stirling Space Engine Program, Tech. Rep. NASA/CR /VOL, Glenn Research Center, August [23] Mason, L. S., A Power Conversion Concept for the Jupiter Icy Moons Orbiter, Tech. Rep. NASA/TM? , Glenn Research Center, September [24] Joseph A. Angelo, J. P. and Buden, D., Space Nuclear Power, Orbit Book Company, Inc,

16 REFERENCES REFERENCES [25] Dawson, J., Advanced Fusion Reactors, Vol. B, Academic Press, New York, 1981, p [26] Cohen, S., A fusion power plant without plasma-material interactions, Tech. Rep. PPPL-3245, Princeton Plasma Physics Laboratory, September [27] Miller, W. O. and Shih, W., Lightweight Carbon-Carbon High-Temperature Space Radiator, [28] Williams, C., Borowski, S., Dudzinski, L., and Juhasz, A., Realizing 2001: A Space Odyssey: Piloted Spherical Torus Nuclear Fusion Propulsion, Tech. Rep. NASA/TM? , NASA Glenn Research Center, March [29] Shea, D. and Morgan, D., The Helium-3 Shortage: Supply, Demand, and Options for Congress, Tech. Rep R41419, Congressional Research Service, Dec [30] Price, H., Hawkins, A., and Radcliffe, T., Austere Human Missions to Mars, AIAA Space 2009 Conference, Pasadena, California, September [31] Bramanti, C., Izzo, D., Samaraee, T., Walker, R., and Fearn, D., Very high delta-v missions to the edge of the solar system and beyond enabled by the dual-stage 4-grid ion thruster concept, Acta Astronautica, Vol. 64, No. 78, 2009, pp [32] Landau, D. and Strange, N., Near-Earth Asteroids Accessible to Human Exploration with High-Power Electric Propulsion, AAS/AIAA Astrodynamics Specialist Conference, Vol. 11, No. 446, [33] Staff, J., Prometheus Project Final Report, Tech. Rep. 982-R120461, Jet Propulsion Laboratory, October [34] Razin, Y., Paluszek, M., Pajer, G., Glasser, A. H., and Cohen, S., Modular Aneutronic Fusion Engine, Space Propulsion 2012, Association Aeronautique et Astronautique de France, Bordeaux, France, May [35] Razin, Y., Paluszek, M., Ham, E., Pajer, G., Mueller, J., Cohen, S., and Glasser, A. H., Compact Aneutronic Fusion Engine, International Astronautical Congress,, Naples, Italy, October [36] Paluszek, M., Hurley, S., Pajer, G., Thomas, S., Mueller, J., Cohen, S., and Welch, D., Modular Aneutronic Fusion Engine for an Alpha Centauri Mission, DARPA 100 Year Starship Conference, Orlando, Florida, September [37] Brophy, J. R., Gershman, R., Strange, N., Landau, D., and Merrill, R. G., 300-kW Solar Electric Propulsion System Configuration for Human Exploration of Near-Earth Asteroids, Tech. rep., Jet Propulsion Laboratory, California Institute of Technology, [38] Chang-Diaz, F., Facts About the VASIMR Engine and its Development, [39] Soulas, G. C., Haag, T. W., Herman, D. A., Huang, W., Kamhawi, H., and Shastry, R., Performance Test Results of the NASA-457M v2 Hall Thruster, Tech. Rep. NASA/TM , NASA Glenn Research Center, Cleveland, Ohio, August [40] Polk, J., Lithium-Fueled Electromagnetic Thrusters for Robotic and Human Exploration Missions, Space Nuclear Conference, July [41] Gardner, J. P. and et al, The James Webb Space Telescope, Space Science Reviews, Vol. 123, 2006, pp [42] Chang, A., Astronauts face radiation threat on long Mars trip, AP, May [43] Gebhardt, C., Delving Deeper into NASA s DSH configurations and support craft, April