Research Papers (2021 – 2026)

A new methodology for achieving the formation and compressional heating of a Field Reversed Configuration (FRC) plasmoid has been investigated both theoretically and experimentally at the University of Washington and MSNW. This approach, based on previous FRC empirical scaling, is expected to achieve fusion gain as large as 10. For the FRC, the fusion gain, G ~ j_p×B_e² where j_p is the FRC poloidal flux and B_e the confining axial magnetic field. G is essentially independent of the FRC radial scale making feasible a small, compact reactor. The necessary j_p (> 60mWb) is achieved by employing a large formation chamber (~ 0.8m radius) combined with the requisite axial magnetic field reversal time (E_q ~ 20kV/m) for generating a high initial temperature (T_i > 1keV) FRC. By employing the dynamic formation procedure, the FRC is accelerated to ~100-150km/s. This subsonic velocity is maintained as the FRC is translated through a series of cylindrical coils of decreasing radius but of sufficient length (> L_FRC) and number (~ 10) to produce essentially an isentropic and adiabatic radial wall-compression of the equilibrium FRC. The final stage is a 12cm diameter, 3-6m long confinement and burn chamber with a vacuum field of 7-9T produced by external solenoidal coils inside the flux conserving cylinder. The vacuum field is compressed by the FRC on insertion to 35 T resulting in fusion gain conditions for the 2-5 ms transit. The large ratio of FRC to inner wall radius (0.85) substantially lowers the FRC edge pressure thereby greatly reducing both particle and thermal losses. The resulting D-T fusion yield for the 3.5MJ FRC is 20-40MJ/pulse. The final ejection and expansion of the FRC into a large, low field mirror chamber provides a mechanism for FRC energy recovery. The experimental justification and the physics basis for the entire process from formation through burn of the Compact FRC Fusion Reactor (CFR2) concept is presented.

The accumulated inventory of spent nuclear fuel in the United States exceeds 90,000 metric tons, with long-term radiological hazards dominated by minor actinides that require isolation for thousands of years. This study details a possible solution to these challenges that also provides significant electrical power. The method employs a fusion system operating in catalyzed deuterium (Cat-D) mode producing the high neutron flux required for this task. It is based on the Compact Field Reversed Configuration Fusion Reactor (CFR2), where the plasma is sequentially compressed to fusion conditions remote from the burn chamber and then translated into the chamber surrounded by the fission blanket. A Monte Carlo neutronics analysis of a subcritical actinide transmutation blanket and the design was optimized. The optimized blanket employs FLiNaK eutectic fluoride salt enriched to 90% ⁶Li and containing 17.53 mol% transuranic fluorides as the combined coolant, moderator, and fissile medium. The design achieves keff=0.9695 with a blanket volume of 13.4 m³. The FRC is pulsed at 4Hz with a tritium fraction of 4%, yielding a neutron flux of 4×10¹⁸ n/s. This fission system burns 683 kg/yr of TRU, produces 9.6 kg/yr of tritium and 1,431 MWth, requiring only 17.5 MWe to power the CFR2.

Preprint – paper accepted for publication in: Nuclear Energy and Technology

Liquid oxygen depots in low Earth orbit could transform space logistics, however every kilogram of stored propellant must itself be launched at a cost of ~22 kg of launch propellant. The Space Gas Station (SGS) removes this penalty by harvesting propellant in situ. A 5 m radius collector captures the ambient atomic oxygen stream at 300 km and stores it as liquid oxygen. An Ambient Ion Accelerator (AIA) comprised of a two-grid electrostatic accelerator exploits the pre-existing ionospheric O⁺ population to provide drag makeup. The AIA requires no pre-ionization stage, neutralizer, or onboard propellant. High gas collection efficiency is achieved with a cryogenic system that avoids the direct collection of the incoming stream by employing a louvered cold-head lining a radiatively cooled conical wall. A 96% capture rate was determined from particle-trajectory calculations with only 7% of the hyperthermal flux reaching the cold-head, thereby keeping required cooling power to 2–4 kW. The reference SGS collects 490 kg of oxygen per year for a launch mass of 170 kg and 17 kW, repaying its launch mass in four months. The analysis identifies the opposed altitude gradients of the neutral and ion populations as the constraint governing orbit selection, motivating an elliptical orbit that collects at perigee and thrusts at apogee. A sample mission analyzes  propellant-neutral removal of a one-tonne spent rocket body from 950 km. Also, a small fleet is shown sufficient to supply the oxidizer for the Mars Ascent Vehicle and the many control thrusters needed for human Mars exploration.

Keywords: in-situ resource utilization; VLEO; atmosphere-breathing electric propulsion; propellant depot; active debris removal; cryogenic collection

Preprint – Submitted to: Acta Astronautica

J Slough – 2024 IEEE Aerospace Conference, 2024 – ieeexplore.ieee.org

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 new 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 as well as limiting the thermal heat load and therefor radiator mass. The FDR concept is based on research conducted at MSNW and the University of Washington on magnetically driven implosions of thin metal bands onto magnetized plasmoids to achieve nuclear fusion. Here, Lithium bands are inductively driven to converge radially and axially forming a thick blanket surrounding the plasmoid. Virtually all of the radiant, neutron and particle energy from the fusion plasma is absorbed by the encapsulating shell thus isolating the thruster and spacecraft from the intense energy release. This energy, in addition to the powerful Ohmic heating in the shell at peak magnetic field compression, is adequate to vaporize and ionize the Lithium. The expansion of this hot, ionized metal propellant is inhibited from blow-back onto the thruster as it is produced inside the large region filled with a divergent axial magnetic field. The propellant thermal energy is consequently converted into high thrust at an optimal Isp through this magnetic nozzle. The back emf from magnetic flux compression due to the expanding Lithium plasma can be utilized to resupply the electrical energy needed for subsequent pulses. This paper outlines these key elements of the FDR concept and the state of their current development. These include: 1) The results from laboratory testing of the metal band compression with sufficient imploded foil kinetic energy (~ 0.5 MJ) to reach conditions required for fusion breakeven conditions. (2) The results from both a 1D code for analyzing the magnetically driven metal band implosion dynamics as well as a full 3D ANSYS® calculation of the implosion dynamics. (3) Based on these results the FDR spacecraft design and mass budget were formulated. (4) A range of mission architectures for which FDR would be enabling, including a rapid, single launch manned Mars mission were developed and are described.

J Slough – 2024 IEEE Aerospace Conference, 2024 – ieeexplore.ieee.org

J Slough – 2023 IEEE Aerospace Conference, 2023 – ieeexplore.ieee.org

To create a more efficient propulsion system for in-space transportation the NEP and NTP systems are now being developed to facilitate the human exploration of Mars. They are, however, unlikely to meet the challenge of deep space without significant advances in efficiency and the ability to use in-situ resources. The concept outlined in this paper is the development of a planetary orbiter spacecraft that can operate on virtually any gas, and is designed in such a way that it can also harvest the planetary atmospheres to obtain the fuel required for the return transit. It is based on experimental results employing a robust plasma creation and sustainment methodology referred to as a Rotating Magnetic Field (RMF) generated Field Reversed Configuration (FRC). In the experiments a Hydrogenic gas was partially ionized and accelerated out of a magnetized cascaded arc source with directed velocities of 35 km/s that is characteristic of the conditions that would confront a Harvesting Planetary Orbiter (HarPO) in the atmospheres of the Jovian planets. The transverse RMF fully ionized the gas, and in the process generated large azimuthal currents which produced an FRC that confined and heated the ions. The stationary, axial guide magnetic field confined the plasma within the chamber wall and conducted the plasma exhaust to form an on-axis plasma jet. In the spacecraft implementation, this system would efficiently process the incoming atmosphere into the necessary thrust and Isp for drag makeup. Based on the ion temperature measurements in the experiments, only a fraction of the gas ingested during the orbital transit would be required as the projected Isp is 12 ksec. The plasma chamber radius, which is the same as the spacecraft intake aperture, need be no larger than that employed in past RMF FRC experiments (0.4m). Given that the HarPO is electrodeless and magnetically insulated, it quite capable of long-term operation on the lighter elements, so that the array of HarPO’s could also be employed as the main thrusters for the transits to and from the outer planets. The analysis and design of a prototype HarPO has been completed based on the drag, power processing, energy and fuel storage requirements that would be encountered in the Jovian atmospheres. The results from this analysis, as well as the architectures for a Neptune and Jupiter mission will be described.

J Slough – 2023 IEEE Aerospace Conference, 2023 – ieeexplore.ieee.org

J Slough 2023 IEEE Aerospace Conference (AERO), 2023•ieeexplore.ieee.org

For deep-space planetary missions, the challenge of orbit insertion at the target planetary body is as critical and potentially as fuel consumptive as launch. The total required velocity change can be on the order of 10s of km/s, resulting in fuel being a large fraction of the spacecraft mass thus limiting valuable payload. Virtually all the planetary bodies in the solar system have an atmosphere of sufficient depth and mass density to provide for rapid vehicle deceleration from aerodrag. Even braking in the Earth’s atmosphere would be highly advantageous for a Martian sample return mission. Employing a solid deflector for orbit insertion has been successful in the past, but due to frictional heating, as well as unknown and changing atmospheric conditions, the braking must be distributed over many elliptical orbital passes into the atmosphere to achieve the required braking. The concept to be elucidated in this paper is one possible solution to the limitations of aerobraking with a solid shield. The method entails the deployment of an arrangement of dipolar magnetic field coils that is populated with plasma and sustained in a manner that provides the necessary impediment to atmospheric flow and thereby creates the desired drag. This braking technique is referred to as a Plasma MagnetoShell (PMS). The requisite drag-inducing interactions are between the incoming neutral’s directed momentum and the magnetically confined ions in the form of charge exchange and elastic scattering. During the interaction, the neutral’s momentum is imparted to the magnetized ion which is then reacted onto the magnetic coils tethered to the spacecraft through field line bending and stretching. It will be shown that there are several unique advantages to PMS braking and aerocapture. Frictional heating is no longer a concern. The aeroshell is now composed of massless magnetic field and plasma. The scale of the magnetic barrier can be as large as needed (~10-100 m) while requiring less than a gram of plasma. While the dipole coil system will be the dominant mass, it is negligible compared to the spacecraft mass. With the ability to control the drag by changing field strength and/or plasma density on a sub-second timescale, the larger drag force and precision trajectory required for aerocapture is now feasible. Since the PMS provides for dynamic control under rapidly changing conditions, it should be suitable for human missions. A laboratory scale PMS was designed, built, and tested in a large vacuum chamber at MSNW. A supersonic flowing neutral jet was aimed a subscale PMS that was mounted on a boom attached to a thrust stand. The tests results validated the concept with the drag force increasing by a factor of 1000X when the PMS was activated. Mission analyses and system designs two missions were also conducted: a Cassini-class Neptune orbiter and a HEOMD-scale Mars orbital insertion. The design included the Plasma MagnetoShell magnet, tether, power processing, and battery mass. The orbital code Copernicus was used to determine optimal trajectories. The results from these studies will also be presented.

J Slough 2023 IEEE Aerospace Conference (AERO), 2023•ieeexplore.ieee.org

An optimal magnetic shielding configuration for significantly reducing astronaut exposure to Galactic Cosmic Radiation (GCR) on long interplanetary missions has been realized, and is referred to as the Dipolar Toroidal Magnetosphere (DTM). This configuration was shown to have the singular ability to deflect the vast majority of the GCR including High Z Energetic (HZE) ions. This external (to the spacecraft) dipolar field is created by an array of unidirectional toroidal High Temperature Superconductor (HTSC) windings mounted externally on the surface of the toroidally-shaped spacecraft habitat. In this way the spacecraft directly supports the magnetic hoop forces generated by the toroidal currents and thereby significantly reduces the structural mass requirements for the shield. The magnitude of the toroidal currents are arranged poloidally to flow so as to maintain the spacecraft shell as a constant flux boundary such that the magnetic flux φ=0 inside the spacecraft keeping the habitat field-free with no need for providing additional coils to cancel the internal fields. As the dipole magnetic field is everywhere perpendicular to the spacecraft habitat, the DTM provides a deflecting shield to all the incoming GCR which is nearly isotropic. In addition, the DTM shields the HTSC coils as well thereby eliminating the secondary particle irradiation hazard, which can dominate over the primary GCR for shields with closed magnetic topologies. With DTM shielding it was found that both the structural and magnet mass as well as the total current requirements were significantly reduced. The DTM shielding effectiveness was calibrated in a large laboratory vacuum chamber experiment using a high energy beam as a surrogate for the GCR encountered in space. MSNW developed 3-D relativistic particle code to evaluate magnetic shielding of GCR and also employed a very accurate commercially available particle trajectory and magnetic field solver, Lorentz-3M, from Integrated Engineering to evaluate shielding effectiveness for the GCR spectrum encountered in space. The codes were employed to evaluate a wide range of magnetic topologies and shielding approaches including prior proposed magnetic shielding concepts. Four criteria for evaluation of the efficacy of magnetic shielding are discussed for three of the most developed concepts including the DTM. They are: (1) Effectiveness of the magnetic shielding using the level of protection achieved for a given magnetic geometry as a function of the magnet mass required. The DTM proved to be an order of magnitude better. (2) Issues with secondaries: The interaction of the GCR with large, exposed magnetic systems has a profound effect on the total radiation exposure inside the habitat independent of the shield effectiveness. (3) Launch and space assembly: The amount of mass, scale of structures, and complexity of assembly. (4) Advantages and other uses for the concept beyond GCR protection such as magneto-aero braking and capture.

There are two critical issues for in-space operations that would be addressed and potentially solved by the concept proposed here. These are: Active Debris Removal (ADR) and in-space maneuverability, e.g. providing mobility options to space assets, as well as in-space repair and reconstruction of disabled assets. There are many challenges associated with in-space maneuverability. Due to the high delta-V requirements, conventional chemical or propellant based systems are implausible, and only a reusable system is capable of keeping costs reasonable. It demands a specialized spacecraft that can maneuver between many target objects, rendezvousing, capturing and transferring assets into new orbits, or deorbiting unwanted space debris. Such an Orbital Transfer Vehicle (OTV) requires the ability to perform hundreds of km/s in delta-V over its mission lifetime. An OTV that can utilize in-situ upper atmospheric resources would enable a host of NASA and DoD missions that require extremely high delta-V in a fast, responsive, and repeatable manner. The Refueling Orbital Navigator (RON) spacecraft accomplishes this in the following way: first it employs a highly elliptical orbit to obtain fuel by scooping up and storing the molecular nitrogen and atomic oxygen encountered during the low altitude perigee periods of its orbit. As RON can thrust at apogee, it can achieve the extensive orbit lowering needed for ADR. Additionally, RON can thrust at perigee to provide drag compensation for low perigee refueling, stable non-Keplerian orbits, or rapid phase changes. In a recent effort at MSNW, funded by NASA’s NIAC program, the design and the initial testing of the RON concept was successfully performed. It was determined that such a revolutionary system involved three key elements. The first was the adoption of Elliptic Collection Orbits (ECOs). In this way the collection of ambient atmospheric gases for propellant could be decoupled from the constant need for drag make-up that has effectively doomed prior efforts at air-breathing propulsion which typically employed circular orbits. The ECO limits the atmospheric drag to a brief period of time at perigee, but allows for solar energy accumulation and storage over the entire orbital period. This, combined with modern lightweight solar panel technology, make up the second critical element by creating a unique opportunity for lightweight, efficient, and high-power electric propulsion (EP) such as the Hall thruster to be employed. Finally, a highly efficient propellant collection system demanded by a refueling OTV can be achieved by a properly designed low-drag scoop geometry coupled to an active collection and storage system that employs turbo-molecular pumping. Such a collection and storage system has been devised and several key elements have been fabricated and successfully tested. Initial orbital dynamics calculations are presented as well as a sample mission analysis for the removal of a metric ton satellite from orbit.

Fission-based power sources have been used in space as early as the Voyager missions. However, the use of nuclear power for manned space exploration has yet to be accomplished. Reactor-based (thermal or electric) propulsion efforts have been stymied by enormous launch costs and extremely long mission times, which endanger the health of the astronauts, and increase the risk of failure due to the need for complex space rendezvous and assembly. All of this is avoided with a compact, high efficiency, single-launch spacecraft such as that projected for the Fission-Powered Pulsed Plasma Propulsion concept (F4P) proposed here. In the F4P scheme, the immense energy of the nuclear reaction is to be harnessed in a more direct and powerful way to propel the spacecraft with the optimal range of jet power (30–100 MW), and at the most efficacious exhaust velocity (30–50 km/sec). It is accomplished in the following manner: (1) Large, thin hoops of Lithium (Li) metal are driven at the proper angle and speed for convergence onto fissile target injected from the F4P thruster throat. (2) The converging Li shells form a thick blanket compressing the fissile material thereby producing a large increase in reactivity and fission energy release. (3) Virtually all of the energy released (plasma thermal, radiant, shell-ohmic, and fission-fragment) is absorbed by the Lithium shell which is subsequently vaporized and ionized. This Li plasma plume expands out against the thruster’s divergent magnetic field which acts as a barrier and directs the plasma flow out of the thruster resulting in minimal interaction with the spacecraft. In this manner, both high exhaust velocity and thrust are achieved with only minor interaction with the spacecraft. To understand the issues involved with this approach, as well as design an experimental test, a 1D analytical model was developed and 3D ANSYS Explicit Dynamics® of the liner implosion were carried out. The inductive coupling and drive efficiency of the compression for a single liner was evaluated with the one-dimensional model which included the key circuit, magnetic field and liner parameters. The model was then used to characterize the liner motion as a function of liner mass, resistivity, stored energy, coil voltage and initial magnetic field. The stability and structural behavior of multiple liners was studied with the 3D ANSYS code. Techniques for controlling liner bucking as well as liner rotation for Rayleigh-Taylor stability were examined with these codes Experimental testing at the relevant scale have validated the metal hoop compression process. The major unknown to be resolved is the functional dependence of the fission energy yield on the several key parameters of the compression process. This must be fully understood and tested to prove the viability of the F4P concept. The current effort is focused on (1) a detailed physics model of the heating and compression of the fissile fuel from a solid to an ionized gas. (2) Calculations to determine the fission amplification and energy yield throughout the compression cycle as a function of fissile mass and density.