An innovative facility for testing high-power, pulsed plasmoid thrusters has been constructed to develop the electrodeless Lorentz force (ELF) thruster concept. It is equipped with a suite of diagnostics optimized to study the physical processes taking place within ELF and evaluate its propulsive utility including magnetic field, neutral gas, and plasma flux diagnostics, a method to determine energy flow into the plasma from the pulsed power systems, and a new type of ballistic pendulum, which enables thrust to be measured without the need for installing the entire propulsion system on a thrust stand. Variable magnetic fields allow controlled studies of plume expansion in a small-scale experiment and dielectric chamber walls reduce electromagnetic influences on plasma behavior and thruster operation. The unique capabilities of this facility enable novel concept development to take place at greatly reduced cost and increased accessibility compared to testing at large user-facilities.
NomenclatureB b , B bias = magnitude of preexisting (vacuum) axial magnetic field B = amplitude of rf rotating magnetic field B z = axial magnetic field change due to RMF driven currents (superscript M denotes maximum) B ext = magnitude of magnetic field external (radially) to the FRC = plasma pressure normalized to external (vacuum) field C = capacitance value = classical skin depth = (2/ 0 ) 1/2 E = electric field vector (subscripts r, , z denote cylindrical components) e = unit of electron charge E k = propellant kinetic energy E k_RMF = kinetic energy derived from electromagnetic input E k_th = kinetic energy derived from conversion of plasma thermal energy E ion = ionization energy E = energy input from Ohmic heating = Hall scaling parameter F = force vector (subscripts r, , z denote cylindrical components) FRC = field reversed configuration = magnetic flux I = current (subscripts r, , z denote cylindrical components) Isp = propellant specific impulse j = current density (subscripts r, , z denote cylindrical components) k = Boltzmann's constant = plasma resistivity e = thruster efficiency = ratio of plasma radius to classical skin depth 0 = magnetic permeability in vacuum m = electron mass ei = electron-ion collision frequency n = plasma density n e = electron density N = electron line density p = plasma pressure = azimuthal cylindrical coordinate Q = circuit quality factor r = radial cylindrical coordinate r p = plasma radius r s = magnetic separatrix radius 1 AIAA Senior Member 2 AIAA Member 3 AIAA Member 2 Figure 1. ELF thruster operating on water vapor. RMF= rotating magnetic field T = plasma temperature T e = electron temperature = RMF pulse length = angular frequency of RMF ce = electron angular frequency in rotating field ci = ion angular frequency in rotating field z = axial cylindrical coordinate
Publiq,reporting burden for this collection of information Is estimated to average 1 hour per response, including the time for revlevning instructions, searching existing data sources gathering and maintaining the data needed, and completing and reviewing this colleofion of information. Send comments regarding this burden estimate or any other aspect of this coliection of information. A national team of researchers Is developing a software system, COLISEUM, which performs calculations of plasma propagation and interaction with arbitrary 3-D surfaces. The applications of COLISEUM are wideranging, but include simulating engine test configurations inside vacuum chambers and predicting sputtering and re-deposltlon on spacecraft surfaces. COLISEUM allows users to easily define complicated 3-D geometries using off the shelf CAD software, then select from a set of plasma expansion models of varying fidelities and numerical complexity to perform the solution. Once the object surfaces are created, the user can run different types of simulations for the same geometry. With this system, low fidelity models can be used to verify the geometry and boundary conditions, and to obtain first-order predictions. Higher fidelity models are then used to obtain more accurate predictions with greater cost in computation time. This paper describes the simulation of a Hall thruster firing inside a vacuum chamber, and the execution of two types of simulations for the same geometry: one in which the equilibrium sputter and deposition rates are calculated, and another In which the Hall thruster plume expansion is calculated. The sputter and deposition calculations are made with an Iterative ray tracing algorithm which takes into account re-sputtering of deposited material. The plume expansion calculations are performed with a particle-ln-cell (PIC) algorithm which Includes wall collisions and wall recombination. Techniques are Incorporated that decrease the computation time required for the PIC simulation to relax to steady state.
Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing this collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB control number. PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. Recent advances in energy storage and solid-state switching enabled the use of peristaltic, pulsed inductive acceleration of nonferritic particles for spacecraft propulsion. Macron Launched Propulsion (MLP) systems electromagnetically accelerate gramsized aluminum particles (i.e. macrons) to achieve exit velocities between 5 and 10km/s, achieving specific impulses between 600 and 1,000s. Research was conducted to analyze this system's potential effects on the orbital debris environment as well as to formulate possible implementations of this technology. Ultimately, the direction, velocity and altitude at which these macrons are fired determine the macron's trajectory and dictate the level of impact upon the orbital debris environment. Research supports the implementation of the technology as a multipurpose orbital maneuvering system but cautions the use of this system in a manner that could result macrons entering into a stable Earth orbital trajectory. 15. SUBJECT TERMS 16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF ABSTRACT 18. NUMBER OF PAGES 19a. NAME OF RESPONSIBLE PERSON
The future of manned space exploration and development of space depends critically on the creation of a dramatically more proficient propulsion system for in-space transportation. This has been recognized for many years, creating a persuasive reason for investigating the applicability of nuclear power in rockets. Nuclear fuel contains energy densities that dwarf the energy of any chemical combustion. However, many such nuclear based propulsion system are not truly feasible as space-based sytems due to size, complexity, cost, or potential hazard. The Fusion Driven Rocket (FDR) described in this work offers a realistic approach to fusion propulsion systems. FDR allows for direct energy transfer to the propellant requiring no conversion to electricity. Addtionally, the propellant requires no significant tankage mass by being a solid, yet can still be rapidly heated and accelerated to high exhaust velocity (> 20 km/s). But perhaps most importantly, unlike many other fusion and fission concepts, there is no significant physical interaction with the spacecraft thereby limiting thermal heat load, spacecraft damage, and radiator mass. This paper will discuss the basic physics of the FDR and the fusion method employed as well as focus on in-depth analysis of the mission architectures enabled by the FDR. Nomenclature α cap , = specific mass of the capacitors α SEP = solar panel specific mass B = magnetic field C = fusion constant, 4.3 x 10 -8 D-T = Deuterium -Tritium ∆V = change in velocity for a mission or transfer ΔT = length of mission or transfer E in = energy input into fusion reaction E out = energy released from the fusion reaction E k = kinetic or propulsive energy f = frequency of operation η T = thrust efficiency ETO = Earth to Orbit FDR = Fusion Driven Rocket FRC = Field Reversed Configuration g 0 = gravitational constant, 9.81 m/s 2 G F = total fusion gain G I = gain from fusion igniton ICF = Interial Confinment Fusion LEO = Low Earth Orbit MIF = Magneto Inertial Fusion M L = liner mass MR = mass ratio M i = initial mass of spacecraft M f = final mass of spacecraft M S = mass of structure M P = mass of propellant M PL = mass of payload P SEP = power from solar panels 2 I sp = specific impulse R = radius of target ρ = density of target SEP = Solar Electric Power V L = linear velocity ϕ ion = Ionization energy, 75 MJ/kg
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