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NASA NTRS2003

Megawatt Electromagnetic Plasma Propulsion

The NASA Glenn Research Center program in megawatt level electric propulsion is centered on electromagnetic acceleration of quasi-neutral plasmas. Specific concepts currently being examined are the Magnetoplasmadynamic (MPD) thruster and the Pulsed Inductive Thruster (PIT). In the case of the MPD thruster, a multifaceted approach of experiments, computational modeling, and systems-level models of self field MPD thrusters is underway. The MPD thruster experimental research consists of a 1-10 MWe, 2 ms pulse-forming-network, a vacuum chamber with two 32 diffusion pumps, and voltage, current, mass flow rate, and thrust stand diagnostics. Current focus is on obtaining repeatable thrust measurements of a Princeton Benchmark type self field thruster operating at 0.5-1 gls of argon. Operation with hydrogen is the ultimate goal to realize the increased efficiency anticipated using the lighter gas. Computational modeling is done using the MACH2 MHD code, which can include real gas effects for propellants of interest to MPD operation. The MACH2 code has been benchmarked against other MPD thruster data, and has been used to create a point design for a 3000 second specific impulse (Isp) MPD thruster. This design is awaiting testing in the experimental facility. For the PIT, a computational investigation using MACH2 has been initiated, with experiments awaiting further funding. Although the calculated results have been found to be sensitive to the initial ionization assumptions, recent results have agreed well with experimental data. Finally, a systems level self-field MPD thruster model has been developed that allows for a mission planner or system designer to input Isp and power level into the model equations and obtain values for efficiency, mass flow rate, and input current and voltage. This model emphasizes algebraic simplicity to allow its incorporation into larger trajectory or system optimization codes. The systems level approach will be extended to the pulsed inductive thruster and other electrodeless thrusters at a future date.

Gilland, James↗

NASA NTRS1980

The applicability of MFD thrusters to satellite power systems

The high power self field MPD thruster uses electromagnetic forces rather than electrostatic to accelerate a neutral plasma. The most attractive application of MPD thrusters to satellite power systems is in the area of electric propulsion for a cargo orbit transfer vehicle (COTV). Calculations were performed in order to compare the performance of a COTV using an ion or MPD propulsion system. Results show that the MPD propulsion system gives a shorter trip time with the same power and payload when compared to the ion thruster propulsion system at either value of specific impulse. More important than the trip time benefit may be the advantage a MPD propulsion system provides in system simplicity. Another interesting COTV concept using MPD thrusters is the use of a remote power supply located on the Earth, at GEO, or somewhere in between to transmit power to the COTV in a microwave transmission. The specific impulse at thrust levels of tens of newtons makes a MPD propulsion system a candidate for stationkeeping and attitude control of large space structures such as a SPS.

Jones, R. M.↗

NASA NTRS1993

Electromagnetic propulsion for spacecraft

Three electromagnetic propulsion technologies, solid propellant pulsed plasma thrusters (PPT), magnetoplasmadynamic (MPD) thrusters, and pulsed inductive thrusters (PIT), were developed for application to auxiliary and primary spacecraft propulsion. Both the PPT and MPD thrusters were flown in space, though only PPT's were used on operational satellites. The performance of operational PPT's is quite poor, providing only approximately 8 percent efficiency at approximately 1000 s specific impulse. However, laboratory PPT's yielding 34 percent efficiency at 2000 s specific impulse were extensively tested, and peak performance levels of 53 percent efficiency at 5170 s specific impulse were demonstrated. MPD thrusters were flown as experiments on the Japanese MS-T4 spacecraft and the Space Shuttle and were qualified for a flight in 1994. The flight MPD thrusters were pulsed, with a peak performance of 22 percent efficiency at 2500 s specific impulse using ammonia propellant. Laboratory MPD thrusters were demonstrated with up to 70 percent efficiency and 700 s specific impulse using lithium propellant. While the PIT thruster has never been flown, recent performance measurements using ammonia and hydrazine propellants are extremely encouraging, reaching 50 percent efficiency for specific impulses between 4000 to 8000 s. The fundamental operating principles, performance measurements, and system level design for the three types of electromagnetic thrusters are reviewed, and available data on flight tests are discussed for the PPT and MPD thrusters.

Myers, Roger M.↗

NASA NTRS1991

Los Alamos NEP research in advanced plasma thrusters

Research was initiated in advanced plasma thrusters that capitalizes on lab capabilities in plasma science and technology. The goal of the program was to examine the scaling issues of magnetoplasmadynamic (MPD) thruster performance in support of NASA's MPD thruster development program. The objective was to address multi-megawatt, large scale, quasi-steady state MPD thruster performance. Results to date include a new quasi-steady state operating regime which was obtained at space exploration initiative relevant power levels, that enables direct coaxial gun-MPD comparisons of thruster physics and performance. The radiative losses are neglible. Operation with an applied axial magnetic field shows the same operational stability and exhaust plume uniformity benefits seen in MPD thrusters. Observed gun impedance is in close agreement with the magnetic Bernoulli model predictions. Spatial and temporal measurements of magnetic field, electric field, plasma density, electron temperature, and ion/neutral energy distribution are underway. Model applications to advanced mission logistics are also underway.

Schoenberg, Kurt↗

NASA NTRS1989

Advanced solar-propelled cargo spacecraft for Mars missions

At the University of Washington, three concepts for an unmanned, solar powered, cargo spacecraft for Mars-support missions have been investigated. These spacecraft are designed to carry a 50,000 kg payload from a low Earth orbit to a low Mars orbit. Each design uses a distinctly different propulsion system: a solar radiation absorption (SRA) system, a solar-pumped laser (SPL) system, and a solar powered mangetoplasmadynamic (MPD) arc system. The SRA directly converts solar energy to thermal energy in the propellant through a novel process developed at the University of Washington. A solar concentrator focuses sunlight into an absorption chamber. A mixture of hydrogen and potassium vapor absorbs the incident radiation and is heated to approximately 3700 K. The hot propellant gas exhausts through a nozzle to produce thrust. The SRA has an I(sub sp) of approximately 1000 sec and produces a thrust of 2940 N using two thrust chambers. In the SPL system, a pair of solar-pumped, multi-megawatt, CO2 lasers in sun-synchronous Earth orbit converts solar energy to laser energy. The laser beams are transmitted to the spacecraft via laser relay satellites. The laser energy heats the hydrogen propellant through a plasma breakdown process in the center of an absorption chamber. Propellant flowing through the chamber, heated by the plasma core, expands through a nozzle to produce thrust. The SPL has an I(sub sp) of 1285 sec and produces a thrust of 1200 N using two thrust chambers. The MPD system uses indium phosphide solar cells to convert sunlight to electricity, which powers the propulsion system. In this system, the argon propellant is ionized and electromagnetically accelerated by a magnetoplasmadynamic arc to produce thrust. The MPD spacecraft has an I(sub sp) of 2490 sec and produces a thrust of 100 N. Various orbital transfer options are examined for these concepts. In the SRA system, the mother ship transfers the payload into a very high Earth orbit and a small auxiliary propulsion system boosts the payload into a Hohmann transfer to Mars. The SPL spacecraft releases the payload as the spacecraft passes by Mars. Both the SRA-powered spacecraft and the SPL-powered spacecraft return to Earth for subsequent missions. The MPD-propelled spacecraft, however, remains at Mars as an orbiting space station. A patched conic approximation was used to determine a heliocentric interplanetary transfer orbit for the MPD propelled spacecraft. All three solar-powered spacecraft use an aerobrake procedure to place the payload into a low Mars parking orbit. The payload delivery times range from 160 days to 873 days (2.39 years).

Auziasdeturenne, J.↗

NASA NTRS2001

High-Power Magnetoplasmadynamic Thruster Being Developed

High-power electromagnetic thrusters have been proposed as primary in-space propulsion options for several of the bold new interplanetary and deep space missions envisioned by the Human Exploration and Development of Space (HEDS) Strategic Enterprise. As the lead center for electric propulsion, the NASA Glenn Research Center is actively involved in the design, development, and testing of high-power electromagnetic technologies to meet these demanding mission requirements. One concept of particular interest is the magnetoplasmadynamic (MPD) thruster, shown schematically in the preceding figure. In its basic form, the MPD thruster consists of a central cathode surrounded by a concentric cylindrical anode. A high-current arc is struck between the anode and cathode, which ionizes and accelerates a gas (plasma) propellant. In the self-field version of the thruster, an azimuthal magnetic field generated by the current returning through the cathode interacts with the radial discharge current flowing through the plasma to produce an axial electromagnetic body force, providing thrust. In applied field-versions of the thruster, a magnetic field coil surrounding the anode is used to provide additional radial and axial magnetic fields that can help stabilize and accelerate the plasma propellant. The following figure shows an experimental megawatt-class MPD thruster developed at Glenn. The MPD thruster is fitted inside a magnetic field coil, which in turn is mounted on a thrust stand supported by thin metal flexures. A calibrated position transducer is used to determine the force provided by the thruster as a function of thrust stand displacement. Power to the thruster is supplied by a 250-kJ capacitor bank, which provides up to 30- MW to the thruster for a period of 2 msec. This short period of time is sufficient to establish thruster performance similar to steady-state operation, and it allows a number of thruster designs to be quickly and economically evaluated. In concert with this experimental research, Glenn is also developing and using advanced numerical simulations to predict the performance of self-field and applied-field MPD thrusters.

LaPointe, Michael R.↗

NASA NTRS1982

Analysis and design of steady transonic flow over airfoils by the method of parametric differentiation

The results reported here are based on applying the method of parametric differentiation (MPD) to transform the nonlinear differential equation governing small-disturbance transonic flow to a linear equation. Implicit approximate factorization and monotone methods were used to accelerate convergence of the linear problem by an order of magnitude over successive line over-relaxation. The relative merits of using MPD are discussed in comparison to conventional small-disturbance applications. Several MPD analyses are performed on an array of airfoils. A design procedure utilizing MPD is discussed and demonstrated for two nonlifting cases.

Halt, D. W.↗

NASA NTRS1988

Use of high temperature superconductors in magnetoplasmadynamic systems

The use of Tesla-class high-temperature superconducting magnets may have an extremely large impact on critical development issues (erosion, heat transfer, and performance) related to magnetoplasmadynamic (MPD) thrusters and also may provide significant benefits in reducing the mass of magnetics used in the power processing system. These potential performance improvements, coupled with additional benefits of high-temperature superconductivity, provide a very strong motivation to develop high-temperature superconductivity (HTS) applied-field MPD thruster propulsion systems. The application of HTS to MPD thruster propulsion systems may produce an enabling technology for these electric propulsion systems. This paper summarizes the impact that HTS may have upon MPD propulsion systems.

Reed, C. B.↗

NASA NTRS1989

Electric Propulsion for Manned Mars Exploration

Advanced high-power electric propulsion systems can significantly enhance piloted Mars missions. An increase in the science payload delivered to Mars and the reduction of the total Earth-departure mass are the major system-level benefits of electric propulsion. Other potential benefits are the return of the cargo vehicle to Earth orbit and the availability of high power in Mars orbit for high-power science and communications. Parametric analyses for sizing the cargo mission vehicle for Mars exploration missions are presented. The nuclear-electric propulsion system thruster size, power level, mass, propellant type and payload mass capability are considered in these system-level trade studies. Descriptions of the propulsion system selection issues for both ion and MPD thruster technologies are also discussed. On a manned Mars mission, the total launch mass for an unmanned cargo vehicle in low earth orbit (LEO) can be reduced by up to 50 percent over the baseline oxygen/hydrogen propulsion system. Because the cargo vehicle is sent to Mars prior to the manned mission, the trip time for the vehicle is not a critical factor. By taking advantage of the high specific impulse (I sub sp) of an ion or a Magneto-Plasma-Dynamic (MPD) thruster system, the total LEO mass is reduced from 590,000 kg for the oxygen/hydrogen propulsion system to 309,000 kg for the MPD system and 295,000 kg for the ion system. Many factors must be analyzed in the design of a electric propulsion Mars cargo vehicle. The propellant selection, the number of thrusters, the power level and the specific impulse are among the most important of the parameters. To fully address the electric propulsion system design, trade studies for the differing ion and MPD propulsion system configurations (thruster power levels, number of thrusters, propellants and power systems) must be conducted.

Palaszewski, Bryan↗

NASA NTRS1989

Advanced Solar-propelled Cargo Spacecraft for Mars Missions

Three concepts for an unmanned, solar powered, cargo spacecraft for Mars support missions were investigated. These spacecraft are designed to carry a 50,000 kg payload from a low Earth orbit to a low Mars orbit. Each design uses a distinctly different propulsion system: A Solar Radiation Absorption (SRA) system, a Solar-Pumped Laser (SPL) system and a solar powered magnetoplasmadynamic (MPD) arc system. The SRA directly converts solar energy to thermal energy in the propellant through a novel process. In the SPL system, a pair of solar-pumped, multi-megawatt, CO2 lasers in sunsynchronous Earth orbit converts solar energy to laser energy. The MPD system used indium phosphide solar cells to convert sunlight to electricity, which powers the propulsion system. Various orbital transfer options are examined for these concepts. In the SRA system, the mother ship transfers the payload into a very high Earth orbit and a small auxiliary propulsion system boosts the payload into a Hohmann transfer to Mars. The SPL spacecraft and the SPL powered spacecraft return to Earth for subsequent missions. The MPD propelled spacecraft, however, remains at Mars as an orbiting space station. A patched conic approximation was used to determine a heliocentric interplanetary transfer orbit for the MPD propelled spacecraft. All three solar-powered spacecraft use an aerobrake procedure to place the payload into a low Mars parking orbit. The payload delivery times range from 160 days to 873 days (2.39 years).

Auziasdeturenne, Jacqueline↗

NASA NTRS1991

Magnetoplasmadynamic Thruster Workshop

On May 16, 1991, the NASA Headquarters Propulsion, Power, and Energy Division and the NASA Lewis Research Center Low Thrust Propulsion Branch hosted a workshop attended by key experts in magnetoplasmadynamic (MPD) thrusters and associated sciences. The scope was limited to high power MPD thrusters suitable for major NASA space exploration missions, and its purpose was to initiate the process of increasing the expectations and prospects for MPD research, primarily by increasing the level of cooperation, interaction, and communication between parties within the MPD community.

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NASA NTRS1991

IRS organigram

Charts and graphs relative to magnetoplasmadynamic (MPD) thruster technology are given. The research activities at the Institute of Space Transportation, University of Stuttgart, are summarized. Information is given on the Institute's Electric Propulsion and Plasma Wind Tunnel; thermal arcjet research; the nozzle-type thruster, DT-IRS; nozzle-type MPD thrusters; a hot anode thruster; the DT6 thruster; the ZT-1 thruster; the cylindrical MPD thruster; and a comparison of continuous and quasi-steady MPD.

Messerschmid, Ernst↗

NASA NTRS1992

Electric propulsion system technology

The work performed in fiscal year (FY) 1991 under the Propulsion Technology Program RTOP (Research and Technology Objectives and Plans) No. (55) 506-42-31 for Low-Thrust Primary and Auxiliary Propulsion technology development is described. The objectives of this work fall under two broad categories. The first of these deals with the development of ion engines for primary propulsion in support of solar system exploration. The second with the advancement of steady-state magnetoplasmadynamic (MPD) thruster technology at 100 kW to multimegawatt input power levels. The major technology issues for ion propulsion are demonstration of adequate engine life at the 5 to 10 kW power level and scaling ion engines to power levels of tens to hundreds of kilowatts. Tests of a new technique in which the decelerator grid of a three-grid ion accelerator system is biased negative of neutralizer common potential in order to collect facility induced charge-exchange ions are described. These tests indicate that this SAND (Screen, Accelerator, Negative Decelerator) configuration may enable long duration ion engine endurance tests to be performed at vacuum chamber pressures an order of magnitude higher than previously possible. The corresponding reduction in pumping speed requirements enables endurance tests of 10 kW class ion engines to be performed within the resources of existing technology programs. The results of a successful 5,000-hr endurance of a xenon hollow cathode operating at an emission current of 25 A are described, as well as the initial tests of hollow cathodes operating on a mixture of argon and 3 percent nitrogen. Work performed on the development of carbon/carbon grids, a multi-orifice hollow cathode, and discharge chamber erosion reduction through the addition of nitrogen are also described. Critical applied-field MPD thruster technical issues remain to be resolved, including demonstration of reliable steady-state operation at input powers of hundreds to thousands of kilowatts, achievement of thruster efficiency and specific impulse levels required for missions of interest, and demonstration of adequate engine life at these input power, efficiency, and specific impulse levels. To address these issues we have designed, built, and tested a 100 kW class, radiation-cooled applied-field MPD thruster and a unique dual-beam thrust stand that enables separate measurements of the applied- and self-field thrust components. We have also initiated the development of cathode thermal and plasma sheath models that will eventually be used to guide the experimental program. In conjunction with the cathode modeling, a new cathode test facility is being constructed. This facility will support the study of cathode thermal behavior and erosion mechanisms, the diagnosis of the near-cathode plasma and the development and endurance testing of new, high-current cathode designs. To facilitate understanding of electrode surface phenomenon, we have implemented a telephoto technique to obtain photographs of the electrodes during engine operation. In order to reduce the background vacuum tank pressure during steady-state engine operation in order to obtain high fidelity anode thermal data, we have developed and are evaluating a gas-dynamic diffuser. A review of experience with alkali metal propellants for MPD thrusters led to the conclusion that alkali metals, particularly lithium, offer the potential for significant engine performance and lifetime improvements. These propellants are also condensible at room temperature, substantially reducing test facility pumping requirements. The most significant systems-level issue is the potential for spacecraft contamination. Subsequent experimental and theoretical efforts should be directed toward verifying the performance and lifetime gains and characterizing the thruster flow field to assess its impact on spacecraft surfaces. Consequently, we have begun the design and development of a new facility to study engine operation with alkali metal propellants.

Brophy, John R.↗

NASA NTRS2018

Acceleration Modes and Transitions in Pulsed Plasma Accelerators

Pulsed plasma accelerators typically operate by storing energy in a capacitor bank and then discharging this energy through a gas, ionizing and accelerating it through the Lorentz body force. Two plasma accelerator types employing this general scheme have typically been studied: the gas-fed pulsed plasma thruster and the quasi-steady magnetoplasmadynamic (MPD) accelerator. The gas-fed pulsed plasma accelerator is generally represented as a completely transient device discharging in approximately 1-10 microseconds. When the capacitor bank is discharged through the gas, a current sheet forms at the breech of the thruster and propagates forward under a j (current density) by B (magnetic field) body force, entraining propellant it encounters. This process is sometimes referred to as detonation-mode acceleration because the current sheet representation approximates that of a strong shock propagating through the gas. Acceleration of the initial current sheet ceases when either the current sheet reaches the end of the device and is ejected or when the current in the circuit reverses, striking a new current sheet at the breech and depriving the initial sheet of additional acceleration. In the quasi-steady MPD accelerator, the pulse is lengthened to approximately 1 millisecond or longer and maintained at an approximately constant level during discharge. The time over which the transient phenomena experienced during startup typically occur is short relative to the overall discharge time, which is now long enough for the plasma to assume a relatively steady-state configuration. The ionized gas flows through a stationary current channel in a manner that is sometimes referred to as the deflagration-mode of operation. The plasma experiences electromagnetic acceleration as it flows through the current channel towards the exit of the device. A device that had a short pulse length but appeared to operate in a plasma acceleration regime different from the gas-fed pulsed plasma accelerators was developed by Cheng, et al. The Coaxial High ENerGy (CHENG) thruster operated on the 10-microseconds timescales of pulsed plasma thrusters, but claimed high thrust density, high efficiency and low electrode erosion rates, which are more consistent with the deflagration mode of acceleration. Separate work on gas-fed pulsed plasma thrusters (PPTs) by Ziemer, et al. identified two separate regimes of performance. The regime at higher mass bits (termed Mode I in that work) possessed relatively constant thrust efficiency (ratio of jet kinetic energy to input electrical energy) as a function of mass bit. In the second regime at very low mass bits (termed Mode II), the efficiency increased with decreasing mass bit. Work by Poehlmann et al. and by Sitaraman and Raja sought to understand the performance of the CHENG thruster and the Mode I / Mode II performance in PPTs by modeling the acceleration using the Hugoniot Relation, with the detonation and deflagration modes representing two distinct sets of solutions to the relevant conservation laws. These works studied the proposal that, depending upon the values of the various controllable parameters, the accelerator would operate in either the detonation or deflagration mode. In the present work, we propose a variation on the explanation for the differences in performance between the various pulsed plasma accelerators. Instead of treating the accelerator as if it were only operating in one mode or the other during a pulse, we model the initial stage of the discharge in all cases as an accelerating current sheet (detonation mode). If the current sheet reaches the exit of the accelerator before the discharge is completed, the acceleration mode transitions to the deflagration mode type found in the quasi-steady MPD thrusters. This modeling method is used to demonstrate that standard gas-fed pulsed plasma accelerators, the CHENG thruster, and the quasi-steady MPD accelerator are variations of the same device, with the overall acceleration of the plasma depending upon the behavior of the plasma discharge during initial transient phase and the relative lengths of the detonation and deflagration modes of operation.

Polzin, Kurt A.↗

NASA NTRS1973

Pulsed electromagnetic acceleration

Direct measurements of the power deposited in the anode of a multimegawatt MPD accelerator using thermocouples attached to a thin shell anode reveal a dramatic decrease in the fractional anode power from 50% at 200 KW input power to less than 10% at 20 MW power. The corresponding local power flux peak at a value of 10,000 W/sq cm at the lip of the anode exhaust orifice, a distribution traced to a corresponding peak in the local current density at the anode. A comparison of voltage-current characteristics and spectral photographs of the MPD discharge using quartz, boron nitride and plexiglas insulators with various mass injection configurations led to the identification of different voltage modes and regions of ablation free operation. The technique of piezoelectric impact pressure measurement in the MPD exhaust flow was refined to account for the effects due to probe yaw angle.

Jahn, R. G.↗

NASA NTRS1974

Measured and Calculated Neutron Spectra and Dose Equivalent Rates at High Altitudes; Relevance to SST Operations and Space Research

Results of the NASA Langley-New York University high-altitude radiation study are presented. Measurements of the absorbed dose rate and of secondary fast neutrons (1 to 10 MeV energy) during the years 1965 to 1971 are used to determine the maximum radiation exposure from galactic and solar cosmic rays of supersonic transport (SST) and subsonic jet occupants. The maximum dose equivalent rates that the SST crews might receive turn out to be 13 to 20 percent of the maximum permissible dose rate (MPD) for radiation workers (5 rem/yr). The exposure of passengers encountering an intense giant-energy solar particle event could exceed the MPD for the general population (0.5 rem/yr), but would be within these permissible limits if in such rare cases the transport descends to subsonic altitude; it is in general less than 12 percent of the MPD. By Monte Carlo calculations of the transport and buildup of nucleons in air for incident proton energies E of 0.02 to 10 GeV, the measured neutron spectra were extrapolated to lower and higher energies and for galactic cosmic rays were found to continue with a relatively high intensity to energies greater than 400 MeV, in a wide altitude range. This condition, together with the measured intensity profiles of fast neutrons, revealed that the biologically important fast and energetic neutrons penetrate deep into the atmosphere and contribute approximately 50 percent of the dose equivalant rates at SST and present subsonic jet altitudes.

Foelsche, T.↗

NASA NTRS1975

Pulsed electromagnetic gas acceleration

Terminal voltage measurements with various cathodes and anodes in a high power, quasi-steady magnetoplasmadynamic (MPD) are discussed. The magnitude of the current at the onset of voltage fluctuations is shown to be an increasing function of cathode area and a weaker decreasing function of anode area. Tests with a fluted cathode indicated that the fluctuations originate in the plasma adjacent to the cathode rather than at the cathode surface. Measurements of radiative output from an optical cavity aligned to examine the current-carrying portion of a two-dimensional, 56 kA magnetoplasmadynamic discharge reveal no lasing in that region, consistent with calculations of electron excitation and resonance radiation trapping. A voltage-swept double probe technique allows single-shot determination of electron temperature and electron number density in the recombining MPD exhaust flow. Current distributions within the cavity of MPD hollow cathodes for various static prefills with no injected mass flow are examined.

Jahn, R. G.↗

NASA NTRS1976

Pulsed thermionic converter study

A nuclear electric propulsion concept using a thermionic reactor inductively coupled to a magnetoplasmadynamic accelerator (MPD arc jet) is described, and the results of preliminary analyses are presented. In this system, the MPD thruster operates intermittently at higher voltages and power levels than the thermionic generating unit. A typical thrust pulse from the MPD arc jet is characterized by power levels of 1 to 4 MWe, a duration of 1 msec, and a duty cycle of approximately 20%. The thermionic generating unit operates continuously but with a lower power level of approximately 0.4 MWe. Energy storage between thrust pulses is provided by building up a large current in an inductor using the output of the thermionic converter array. Periodically, the charging current is interrupted, and the energy stored in the magnetic field of the inductor is utilized for a short duration thrust pulse. The results of the preliminary analysis show that a coupling effectiveness of approximately 85 to 90% is feasible for a nominal 400 KWe system with an inductive unit suitable for a flight vehicle.

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