Title of Grant / Cooperative Agreement: The Fusion Driven Rocket: Nuclear Propulsion through Direct Conversion of Fusion Energy Type of Report: Summary of Research Name of Principal Investigator: John Slough Period Covered by Report: 9/10/2012 - 3/10/2015 Name and Address of recipient's institution: 8551 154th Avenue NE, Redmond, WA 98052 NASA Grant / Cooperative Agreement Number: NNX12AQ61G Reference 14 CFR § 1260.28 Patent Rights Have any Subject Inventions / New Technology Items resulted from work performed under this Grant / Cooperative Agreement? No (marked) Details can be provided in the body of the Summary of Research report. Reference 14 CFR § 1260.27 Equipment and Other Property Is there any Federally Owned Property, either Government Furnished or Grantee Acquired, in the custody of the Recipient? No (marked) Reference 14 CFR § 1260.22 Technical publications and reports (December 2003) The future of manned space exploration and development of space depends critically on the creation of a dramatically more efficient propulsion architecture for in-space transportation. A very persuasive reason for investigating the applicability of nuclear power in rockets is the vast energy density gain of nuclear fuel when compared to chemical combustion energy. The Fusion Driven rocket (FDR) represents a revolutionary approach to fusion propulsion where the power source releases its energy directly into the propellant, not requiring conversion to electricity. It employs a solid lithium propellant that requires no significant tankage mass. The propellant is rapidly heated and accelerated to high exhaust velocity (> 30 km/s), while having no substantial physical interaction with the spacecraft thereby avoiding damage to the rocket and limiting both the thermal heat load and radiator mass. The key to achieving this stems from research at MSNW and the UW on the magnetically driven implosion of metal foils onto a magnetized plasma target to obtain fusion conditions. A logical extension of this work leads to a method that utilizes these metal shells (or liners) to not only achieve fusion conditions, but to serve as the propellant as well. Several low-mass, magnetically-driven metal liners are inductively driven to converge radially and axially and form a thick blanket surrounding the target plasmoid and compress the plasmoid to fusion conditions. Virtually all of the radiant, neutron and particle energy from the plasma is absorbed by the encapsulating, thick metal blanket thereby isolating the spacecraft from the fusion. This energy, in addition to the intense Ohmic heating at peak magnetic field compression, is adequate to vaporize and ionize the metal blanket. The expansion of this hot, ionized metal propellant through a magnetically insulated nozzle produces high thrust at the optimal Isp. The energy from the fusion process, along with the waste heat, is thus utilized at very high efficiency. The basic scheme for FDR is illustrated and described in the report (see Fig. 2). The two most critical issues in meeting challenges introduced employing magneto-inertial fusion as the power source is driver efficiency and "stand-off" - the ability to isolate and protect fusion and thruster from the resultant fusion energy. By employing metal shells for compression, it is possible to produce the desired convergent motion inductively by inserting the metal sheets along the inner surface of cylindrical or conically tapered coils. Both stand-off and energy efficiency issues are solved by this arrangement. This two year effort focused on achieving three key criteria for the Fusion Driven Rocket to move forward for technological development: (1) the physics of the FDR must be fully understood and validated, (2) the design and technology development for the FDR required for its implementation in space must be fully characterized, and (3) an in-depth analysis of the rocket design and spacecraft integration as well as mission architectures enabled by the FDR need to be performed. A subscale, laboratory liner compression test facility was assembled at the University of Washington Plasma Dynamics Laboratory with sufficient liner kinetic energy (~ 0.5 MJ) to reach conditions required for fusion breakeven conditions. Detailed experimental studies of the dynamic behavior of the driven liners as well as liner convergence and magnetic compression were performed. The development of both the 1D liner dynamics code and the full 3D ANSYS® liner calculations was achieved. The characterization of both the FDR and spacecraft as well as a design architecture analysis was conducted that included an examination of a wide range of mission architectures and destinations for which this fusion propulsion system would be enabling or critical. In particular a rapid, single launch manned Mars mission was developed. A = cross sectional area of liner As = surface area of liner β = ratio of plasma pressure to magnetic pressure B = magnetic field Be = external magnetic field Bin = internal magnetic field B0 = internal field at peak compression C = capacitance δ = liner thickness δgap = initial gap between driver coil and liner ε = FRC elongation ratio (ls/2rs) EB = magnetic field energy Efus = fusion energy EL = liner kinetic energy F = force on liner FRC = field reversed configuration gAl = action constant for aluminum gM = action constant for material m G = fusion gain I = current Isp = specific Impulse L = inductance Lgap = inductance of coil-liner gap l0 = FRC length at peak compression ls = FRC length ML = liner mass n0 = plasma density at peak compression μ0 = free space permeability PB = magnetic pressure ρm = density of material m r = radius variable rc = coil radius r0 = liner radius at peak compression rs = FRC separatrix radius R = object radius rL = liner radius Δr = liner radial displacement from magnetic force ⟨σv⟩ = D-T fusion cross section dt = time step Te = electron temperature in eV Ti = ion temperature in eV T0 = plasma total temperature at peak compression τD = liner dwell time vL = liner velocity vmax = maximum liner velocity VC = initial voltage on capacitors w = liner width The future of manned space exploration and development of space depends critically on the creation of a dramatically more efficient propulsion system for in-space transportation. This has been recognized for many years. Not as well recognized is the need for high power, high Isp propulsion in Low Earth Orbit (LEO) as it relates to several unmanned orbital maneuvers that are quite costly or not feasible with current chemical propulsion rockets. Of particular value to both military and commercial interests would be a high power orbital tug that can shuttle numerous payloads of several metric tons (MT) from LEO to GEO. A very persuasive reason for investigating the applicability of nuclear power for these missions is the vast energy density gain of nuclear fuel when compared to chemical combustion energy. The combustion of hydrogen and oxygen has an energy release of 13 MJ/kg, whereas the fission of 235U yields approximately 8 x 10^7 MJ/kg and the fusion of deuterium and tritium has a 3.6 x 10^8 MJ/kg yield. So far, the use of fission energy represents the nearest term application of nuclear power for propulsion. Several fission based propulsion schemes have been proposed for in-space transportation, including pulsed nuclear explosions and the Nuclear Thermal Rocket (NTR). In the NTR a cooling fluid or propellant is passed through a core of material that has been heated by fission. This makes the NTR effectively a heated gas rocket. With the present limitations of materials, NTR gas temperatures cannot exceed chemical propulsion gas temperatures. The use of hydrogen provides for an increase in Isp to 900 s. With ΔV ~ 9 km/sec the propellant mass is reduced by an order of magnitude for a given spacecraft mass. Unfortunately, this is considerably offset by increased spacecraft mass (payload, structure, shielding, tankage etc.). A significant mass is required for the low mass density propellant (H2) as the specific gravity of liquid hydrogen is around 0.07, compared to 0.95 for an O2-H2 chemical engine. The net result then is a propulsion system that is better than chemical, but not enough to really be a "game changer". Proposed Nuclear Electric Propulsion (NEP) systems employ high Isp thrusters like the ion and Hall thrusters which solves the propellant Isp issue. The problem for NEP is the inherent inefficiency of the electrical power generation. Shedding the excess reactor heat requires an enormous radiator mass. The large reactor and power conversion masses just add to this problem making for too low a specific power (ratio of jet power to system mass) for rapid space transport. Invoking nuclear fusion for space propulsion, at least as it has been envisioned up till now, does not significantly alter this picture as it has been developed primarily as an alternate source for electrical grid power. This endeavor is far from completion, and even if nuclear fusion were to be eventually developed for terrestrial power generation, the resulting power plant would be unlikely to have any role in space propulsion for all the same reasons that trouble NEP - but worse. Past efforts in this regard have all come to be dismissed, and rightfully so, primarily for the following two reasons. The first has been alluded to already. The propulsion system is reactor based. The straightforward application of a reactor-based fusion-electric system creates a colossal mass and heat rejection problem for space application. In a detailed analysis for the most compact tokamak concept, the spherical torus, spacecraft masses of 4000 metric tons (mt) were projected. The maximum launch mass would need to be less than 150 MT if current chemical rockets are used for launch to LEO. The second is the use of the fusion plasma and/or the fusion products as propellant. Due to the resultant high exhaust velocities, the enormous mismatch from the appropriate exhaust velocity leaves the spacecraft with either insufficient thrust or enormous power requirements. Exotic fuels (3He) or fuels with marginal fusion reaction efficiencies (e.g. P-11B) with large circulating powers are also required. If one were to imagine the optimal solution to this predicament, it would be a fusion propulsion system where (1) the power source releases its energy directly into the propellant, not requiring conversion to electricity. (2) It would employ a solid propellant that would require no significant tankage mass. (3) The propellant is readily heated and accelerated to high exhaust velocity (> 20 km/s), and where the propellant (4) has no significant physical interaction with the spacecraft thereby limiting thermal heat load, spacecraft damage, and radiator mass. The Fusion Driven rocket to be detailed in this paper represents an attempt to incorporate all four of these attributes, and in so doing, lead to a practical propulsion system that is capable of taking full advantage of the promise that fusion energy represents. In this section the physics and method for achieving the compressional heating required to reach fusion gain conditions based on the compression of a Field Reversed Configuration plasmoid (FRC) is elucidated. In brief, an inductive technique is employed to accelerate an array of thin, predominantly lithium metal bands radially inward to create a three dimensional compression of the target FRC in a manner compatible with application in space. A. Background As first demonstrated at Trinity site (fission) and then at Eniwetok Atoll (fusion), the ignition of nuclear fuels have certainly confirmed the ability to produce copious energy yields from nuclear energy, dwarfing that of the Atlas V by many orders of magnitude. The challenge is how to have the release of nuclear energy occur in such a manner as to be a suitable match to that desired for manned planetary transportation or high power orbital maneuvers: multi-megawatt jet power, low specific mass (~ 1 kg/kW) at high Isp (> 2,000 s). It would appear that for at least nuclear fission, there is no real possibility of scaling down to an appropriately low yield as a certain critical mass (scale) is required to achieve the supercritical chain reaction needed for high energy gain. Fission nuclear pulse propulsion then, such as that envisioned in the Orion project, ends up with a thrust in the millions of megatons which would only be suitable for spacecraft on the order of 107 million MT - the mass of over 100 aircraft carriers! Fortunately, the critical mass/scale for fusion ignition can be much smaller. The criteria to achieve D-T fusion ignition, at a nominal fuel (plasma) temperature of 10 keV, is the attainment of a density-radius product of ρ·R ~ 0.1 g/cm^2. This can be accomplished with a three dimensional compression of a spherical cryogenic fuel pellet of millimeter scale. Here it is assumed that the inertia of the small pellet is sufficient to confine the plasma long enough for the burn to propagate through the pellet and thereby produce an energy gain G ~ 200 or more (G = fusion energy/initial plasma energy). This Inertial Confinement Fusion (ICF) approach has been actively pursued for decades as it represents essentially a nano-scale version of a fusion explosive device. Because of the small scale and tiny masses, the energy delivery system required to heat the pellet to fusion temperature must be capable of doing so on the nanosecond timescale. It appeared that the most promising solution to accomplish this is with a large array of high power pulsed lasers focused down on to the D-T pellet. While the anticipated energy yield is in the range appropriate for propulsion (E ~ 20-100 MJ), the scale and mass of the driver (lasers and power supplies) is not, as it requires an aerial photograph to image the full system. There have been however three breakthrough realizations in the last several years that have provided the keys to achieving inertial fusion at the right scale in an efficient and appropriate manner for space propulsion. They primarily concern the enhanced confinement provided by significant magnetization of the target plasma which considerably eases the compressive requirements for achieving fusion gain and even fusion ignition. This new approach to fusion is aptly referred to as Magneto-Inertial Fusion (MIF), and will now be briefly described. B. Magneto Inertial Fusion The notion of using means other than an array of high power lasers to compress the target to fusion conditions goes back as far as the 1950's. Heavy ions and metal shells (liners) were two of the most promising. They all had in common the basic approach of ICF, i.e. the outer shell or liner is driven directly or indirectly inward compressing the inner target to fusion conditions. Regardless of method, this compression must be uniform, intense and accomplished with great precision resulting in large, high voltage and expensive driver systems. By the mid-nineties it was realized that the presence of a large magnetic field in the target would substantially suppress the thermal transport, and thus lower the imploding power needed to compress the target to fusion conditions. With more time before the target plasma thermal energy was dissipated, a much more massive confining shell could be employed for direct compression, with the dwell time of the confining (metal) shell now providing for a much longer fusion burn time. The liner does not need to be propelled inward by ablation but could be driven by explosives or even magnetic fields. In a seminal paper by Drake et al. it was shown that with a fully three dimensional imploding shell on to the magnetized target, fusion gain could be achieved on a small scale with sub-megajoule liner (shell) kinetic energy. There was no known way to accomplish this at that time, but it was feasible at least in theory. The second major theoretical result was obtained by Basko et al. who showed that for a sufficiently magnetized target plasma, fusion ignition would occur even when the restrictive condition that ρ·R > 0.1 g/cm^2 was far from being met. Ignition was now possible as long as the magnetic field-radius product, B·R > 60 T-cm. Thus fusion ignition could be obtained for MIF targets with much lower compression than required for ICF as Fig. 1 indicates. The final critical element to enable fusion energy to be utilized for space propulsion was a practical method to directly channel the fusion energy into thrust at the appropriate Isp. It is believed that such a method has been determined that is supported by both theory and experiment. It was clear that fusion ignition conditions could be achieved at small scale by transferring the kinetic energy of a significantly more massive metal shell than the target plasma to compress it to high density and temperature. The question now becomes: (1) how to do this without invoking a massive and complex driver (2) how to do it in a manner that is efficient and capable of repetitive operation (3) how to create a suitable magnetized plasma target, and (4) how to transfer the fusion energy into a suitably directed propellant at optimal exhaust velocities and powers for the missions mentioned earlier. The key to answering all four "hows" is based on previous experimental work employing inductively driven liners to obtain megabar magnetic pressures, and magnetic compression of Field Reversed Configuration (FRC) plasmoids to fusion conditions. A logical extension based on these results leads to a propulsion method that utilizes these metal shells to not only achieve fusion conditions, but then to become the propellant as well. The basic scheme for Fusion Driven Rocket (FDR) is illustrated and described in Fig. 2. The two most critical matters in meeting challenges (1) and (2) for MIF, and all ICF concepts for that matter, is driver efficiency and "stand-off" - the ability to isolate and protect driver and thruster from the resultant fusion energy release. By employing metal shells for compression, it is possible to produce the desired convergent motion inductively by inserting the metal shells along the inner surface of cylindrical or conically tapered coils. Both stand-off and energy efficiency issues are solved by this arrangement. The metal shell can be positioned up to a meter or more from the target implosion site with the coil driver both physically and electrically isolated from the shell. The driver efficiency can be quite high as the coil driver is typically the inductive element of a simple oscillating circuit where resistive circuit losses are a small fraction of the energy transferred. Even though there is essentially no magnetic field within the liners initially, there is enough flux leakage during the inward acceleration that at peak compression the axial magnetic field that is trapped inside the now greatly thickened wall can reach as high as 600 T. As will be seen this field is considerably higher than that required for the compression of an FRC to achieve ignition and substantial fusion gain. Spaced-based fusion demands a much lower system mass. The lowest mass system by which fusion can be achieved, and the one to be employed here, is based on the very compact, high energy density FRC (see Fig. 3). It is of paramount advantage to employ a closed field line plasma that has intrinsically high β (plasma/magnetic pressure ratio), and that can be readily translated and compressed, for the primary target plasma for MIF. Of all fusion reactor embodiments, only the FRC plasmoid has the linear geometry, and sufficient closed field confinement required for MIF fusion at high energy density. Most importantly, the FRC has already demonstrated both translatability over large distances as well as the confinement scaling with size and density required to assure sufficient lifetime to survive the timescale required for compression. FRCs can, and have been generated with enough internal flux to easily satisfy the B·R ignition criteria at peak compression. At a nominal liner converging speed of 3 km/s, a 0.2 m radius FRC typical of operation on the LSX FRC device would be fully compressed in 67 μs which is only a fraction of the lifetime that was observed for these FRCs (~ 1 msec). Finally, to complete the fourth challenge, a straightforward way to convert the fusion energy into propulsive energy must be devised. It is accomplished by employing an inductively driven thin metal liner to compress the magnetized plasma. As the radial and axial compression proceeds, this liner coalesces to form a thick (r > 5 cm) shell that acts as a fusion blanket that absorbs a large fraction of the fusion energy as well as the radiated plasma energy during the brief fusion burn time. This superheated blanket material is subsequently ionized and rapidly expands inside the divergent magnetic field of the nozzle. Here the thermal energy of the plasma is converted into directed propulsive thrust much like in a conventional nozzle but with the crucial difference being that the isolation provided by the magnetic field protects the chamber wall from bombardment by the energetic plasma ions. It would be possible to also derive the electrical energy required for the driver system from the back emf experienced by the conical magnetic field coil circuit via flux compression. It was found however that the power required for recharging the energy storage modules for the metal liner driver coils could readily be obtained from conventional solar electric power. For very rapid, high power missions, the flux compressor/generator option could be adopted. For a near-term manned mars mission, solar electric requires the least technology development, lowest cost, and is already at the highest TRL level. The analysis of the liner implosion was carried out for both the subscale validation experiment now being performed at the University of Washington (UW), as well as what would be required for a full scale FDR prototype. The dynamics of the liner implosion are governed by: ML (d^2 r / dt^2) = - ((Be^2 / 2μ0) - (Bin^2 / 2μ0)) 2π r w, (1) where ML is the liner mass, and w the liner width. During rapid liner acceleration very little flux diffuses through the liner (Bin << Be). In a variant of the virial theorem, maximum transfer of stored energy Es (= 1/2 C Vc^2) into liner kinetic energy occurs when half is expended in driving magnetic field energy EB. For compression field over short distance Δr: EB = (Be^2 / 2μ0) 2π rL Δr w = 1/4 C Vc^2 => Be = 7.8 x 10^-4 (C / w)^(1/2) (Vc / rL), (2) Equating remaining stored energy with liner kinetic energy: vL = (C / (π rL w δ ρm))^(1/2) (Vc / 2), (3) where ML = 2π rL w δ ρm with δ being thickness and ρm density. At peak compression: EL = 1/2 ML vL^2 = 3 n0 k T0 (4/3 π r0^3 ε) = (B0^2 / μ0) π r0^3 ε, (4) Fusion energy produced during dwell time τD: Efus = 1.2 x 10^-12 n0^2 ⟨σv⟩ (4/3 π r0^3 ε τD) = 1.1 x 10^-42 n0^2 T0^2 (r0^4 / vL) ε, (5) Fusion gain: G = Efus / EL = 1.73 x 10^-3 (ML / l0)^(1/2) B0 = 4.3 x 10^-8 sqrt(ML) EL^(11/8), (6) where l0 = 2 r0 ε. Adiabatic scaling laws yield: EL ~ B0^2 r0^2 l0 ~ B0^(4/5) and l0 ~ r0^(2/5) ~ B0^(-1/5), (7) Energy per unit liner mass: Efus / ML = G (EL / ML) = 4.3 x 10^-8 ML^(15/8) vL^4.75, (8) Current integral for vaporization limit: ∫0^tm I^2 dt = gM A^2, (9) vmax = 6.8 x 10^10 (gM / gAl) (δ / ρm), (10) For lithium, vmax (km/s) = 16 δ (mm). For validation, aluminum is used. A manned Mars case study was chosen as the reference mission. NASA Design Reference Architecture (DRA) 5.0 in 2009 showed that chemical propulsion requires up to 12 heavy lift launches and 1,252 MT in LEO, while Nuclear Thermal Propulsion (NTP) requires 9 launches and 849 MT in LEO, with mission durations of 914 to 1680 days. Table 1. Mission Architecture Goals: - 90 day transit times to and from Mars - Single launch to Mars (130 MT IMLEO) - No pre-deployed assets - 22 MT Payload mass - Full propulsive MOI - Full propulsive EOI - Reusable spacecraft Table 2. Spacecraft Dry Mass Summary: - Mars Entry Vehicle: 22 MT - Crew/Consumables: 2 MT - Transit habitat (MTV): 15 MT - FDR spacecraft: 15 MT - Total Dry Mass: 54 MT Table 3. Mission Assumptions: - Payload mass: 63 MT - Spacecraft mass: 15 MT - Earth Orbital Altitude: 407 km - Mars park orbit: 1 sol (250 x 33,793 km) - Capacitor specific mass: 2.5 J/g - Solar panel specific mass: 200 W/kg - Tankage fraction: 10% - Isp: 5000 s - Fusion Gain: 200 Table 4. Mission Maneuvers: - TMI: ΔV = 7.6 km/s, ΔT = 6.5 days, Mi = 116.4 MT, Mf = 99.6 MT, Mp = 16.8 MT - MOI: ΔV = 13.8 km/s, ΔT = 9.5 days, Mi = 99.6 MT, Mf = 75.1 MT, Mp = 24.5 MT - TEI: ΔV = 16.2 km/s, ΔT = 5.8 days, Mi = 53.1 MT, Mf = 38.2 MT, Mp = 14.9 MT - EOI: ΔV = 11.8 km/s, ΔT = 3.2 days, Mi = 38.2 MT, Mf = 30.0 MT, Mp = 8.2 MT Table 5. Mission Launch Requirements Comparison: - Number of launches: FDR = 1, DRA 5 NTP = 9, DRA 5 chemical = 12 - IMLEO (MT): FDR = 134, DRA 5 NTP = 849, DRA 5 chemical = 1252 - Mission time (days): FDR = 210, DRA 5 NTP = 914(1680), DRA 5 chemical = 914(1680) Table 6. FDR Spacecraft Mass Budget: - Spacecraft structure: 3.4 MT - Lithium containment vessel: 0.1 MT - FRC formation system: 0.5 MT - Propellant Feed mechanism: 1.2 MT - Energy storage: 2.5 MT - Liner driver coils: 0.3 MT - Energy delivery system: 1.2 MT - Solar Panels: 2.0 MT - Thermal management: 1.1 MT - Magnetic expander: 0.2 MT - Margin: 2.5 MT - Spacecraft Mass: 15 MT - Crew habitat (DRA5.0): 63 MT - Lithium Propellant: 56 MT - Total Mass: 134 MT The spacecraft is designed for < 15 MT dry mass and produces 36 MW of jet power at Isp = 5000 s. Energy storage requires a 6.2 MJ pulse capacitor bank with 2.5 J/g specific energy (~2.5 MT mass), accompanied by a 1.2 MT energy delivery system. Power is supplied by 180 kW solar panels scaled to 2.0 MT (at 200 W/kg Earth irradiance). The magnetic driver coils contribute 0.5 MT, thermal management radiators reject 360 kW (1.1 MT), FRC formation hardware is 500 kg, lithium propellant storage is 100 kg plus 1.2 MT feed mechanism. The total spacecraft length is 34-45 m with a 5 m diameter saddle truss containing seven 9 m-long spools of lithium propellant (68 MT total). Experimental studies at the UW Plasma Dynamics Laboratory utilized subscale aluminum liner implosions. Initial tests with 0.41 m radius liners driven by a 392 kJ capacitor bank demonstrated magnetic compression and validated 3D ANSYS Explicit Dynamics Solver (AEDS) modeling. Follow-on experiments used two smaller conical/cylindrical liners (10.3 cm radius, 5 cm width, 0.4 mm thickness) to demonstrate 3D convergent compression with peak magnetic fields exceeding 100 T. Symmetrical folding/pleating and rotation (via rotating quadrupole magnetic field, RMF) were analyzed to eliminate buckling and Rayleigh-Taylor instabilities. Scaling laws for FRC confinement: τN = 3.2 x 10^-15 ε^(1/2) xs^0.8 rs^2.1 n^0.6, (18) Lawson triple product for 15 keV compressed plasma: n τ Ti ~ (6.3 x 10^24)(7.5 x 10^-6)(15) = 7.1 x 10^20 keV-m^-3-s, (19) Rotational frequency required for RT stabilization: Ω0^2 >= (2 / (5 μ0 ρAl ri^4 ln(ri / r0))) r0^2 B0^2 => Ω0 >= 146 rad/sec (Ωi >= 3.7 rad/sec). (24) Parametric studies of a parallel half-turn coil geometry (15 cm inner radius, 3-inch turn width, 20 mil half-turn gap, 3 cm coil spacing) were conducted using a custom MATLAB 1D liner code coupled with ANSYS Maxwell 3D and Moqui. Key parameter sweep findings: - Liner thickness: 20 to 30 mil provides optimal kinetic energy efficiency and velocity while avoiding excess Ohmic heating and premature melting. - Stray inductance: Lowering stray inductance dramatically improves bank-to-kinetic energy conversion efficiency (up to ~60%). - Bank voltage: 25 kV to 30 kV offers the best trade-off for kinetic energy conversion and collapse time. - Capacitance: Values above 5500-6000 μF yield diminishing returns. - Initial bias field: Varying bias field from -1.3 T to 0 T does not penalize efficiency, allowing flexible FRC injection matching. The 1D code utilizes a first-order Euler solver with Lagrangian time-stepping, accounting for dynamic circuit equations, static field calculations, magnetic field breed-through, temperature-dependent resistivity, Ohmic heating, phase change latent heat, and dynamic B-field corrections. 1. Report Date: 08/06/2015 2. Report Type: Final 3. Dates Covered: 9/10/2012 - 3/10/2015 4. Title: The Fusion Driven Rocket: Nuclear Propulsion through Direct Conversion of Fusion Energy 5b. Grant Number: NNX12AQ61G 6. Authors: John Slough, Anthony Pancotti, and Akihisa Shimazu 7. Performing Organization: MSNW LLC, 8551 154th Avenue NE, Redmond, WA 98052 8. Performing Organization Report Number: FDR - Final 9. Sponsoring Agency: NASA Headquarters, 300 E Street Southwest, Washington, DC 20546 12. Distribution/Availability Statement: A - Approved for Public Release 14. Abstract: This two year effort focused on achieving three key criteria for the Fusion Driven Rocket to move forward for technological development: (1) the physics of the FDR must be understood and validated, (2) the design and technology development for the FDR required for its implementation in space must be fully characterized, and (3) an in-depth analysis of the rocket design and spacecraft integration as well as mission architectures enabled by the FDR need to be performed. All criteria were successfully completed and the results are detailed in this report. 15. Subject Terms: Fusion Driven Rocket, plasma propulsion, rocket design, mission analysis 18. Number of Pages: 115 19a. Name of Responsible Person: John Slough (525) 867-8900