Nuclear Electric Propulsion Research Articles

Using magnesium hydride as the future dual-mode propellant for spacecraft: A simulation investigation

A dual-mode propulsion system stands as the important viable pathway for future multi-purpose deep space missions by the spacecraft, encompassing both nuclear thermal propulsion and nuclear electric propulsion. H2 and Mg emerge as potential mediums for nuclear thermal propulsion and nuclear electric propulsion, respectively, with the possibility of utilizing a single material, magnesium hydride (MgH2), to serve both purposes. However, how to release hydrogen completely and effectively in high hydrogen desorption pressure while retaining magnesium is the key problem. In this work, we proposed a novel MgH2 storage tank for zero gravity environments and investigated its hydrogen desorption process at high pressures of 4 ∼ 8 MPa using simulation, which is crucial for the hydrogen supply of nuclear thermal propulsion. We examined the effects of hydrogen desorption pressure, inlet temperature, and flow velocity of argon (Ar) on the desorption process. The simulation results indicated complete decomposition of MgH2 can be realized under 4 ∼ 8 MPa at 873 K. Moreover, the desorption pressure, inlet temperatures, and high flow velocities of Ar are suggested as 4 MPa, 773 ∼ 785 K, and 45 ∼ 100 m s−1 to achieve an average hydrogen desorption rate per unit thermal power of ≥ 20 mgH2 s−1. This work marks the first confirmation that MgH2 can decompose into H2 and Mg under high hydrogen desorption pressure at suitable thermal power, paving the way for its application as the hydrogen and Mg storage media in the dual-mode propulsion system of spacecraft.

Comparison of Propulsion Options for Interplanetary Missions

This paper investigates propulsion options and trajectory designs for a manned mission from Earth to Mars, followed by a mission to Ceres. It evaluates various propulsion technologies, including traditional chemical propulsion, nuclear thermal propulsion (NTP), and electric propulsion systems such as ion thrusters and Hall effect thrusters. The analysis highlights the limitations of chemical propulsion in terms of energy efficiency and payload capacity while underscoring the potential of NTP to reduce travel time and improve crew safety due to its higher specific impulse. The study also examines the feasibility of solar and nuclear electric propulsion, focusing on their high efficiency and suitability for extended missions, and considers innovative concepts like solar sails and fusion propulsion for their theoretical advantages in thrust and velocity. Low-thrust trajectories are analyzed for their ability to minimize overall delta-V requirements, which is essential for mission success. For the Mars mission, chemical propulsion systems currently under development are evaluated alongside electric propulsion powered by nuclear reactors, which could significantly reduce travel times and propellant needs. The additional challenges of propelling a spacecraft to Ceres, a more distant destination in the asteroid belt, shift the focus to electric propulsion options, particularly advanced nuclear-electric systems, which offer the potential to enable human exploration within a reasonable timeframe. The paper concludes that the optimal propulsion system for a manned mission to Mars and Ceres requires balancing travel time, technical feasibility, and crew safety, emphasizing the necessity for further development of advanced electric propulsion technologies to facilitate ambitious human deep-space exploration.

Spacecraft Medium Voltage Direct-Current (MVDC) Power and Propulsion System

This paper introduces a medium voltage direct-current (MVDC) system for large spacecraft megawatt-scale (MW) power and propulsion systems intended for interplanetary transport, including missions to the Moon and Mars. The proposed MVDC system includes: (i) A nuclear electric propulsion (NEP) that powers a permanent magnet (PM) generator whose output is rectified and connected to the MVDC bus. (ii) A solar photovoltaic (PV) source that is interfaced to the MVDC bus using a unidirectional boost DC-DC converter. (iii) A backup battery energy storage system (BESS) that connects to the MVDC bus using a bidirectional DC-DC boost converter. (iv) A dual active bridge (DAB) converter that controls the power to the spacecraft’s electric thruster. The NEP serves as the main power source for the spacecraft’s electric thruster, while the solar PV and BESS are intended to provide power for the payload and spacecraft’s low-voltage power system. The paper will (i) provide a review of the spacecraft MVDC power and prolusion system highlighting state-of-the-art main components, (ii) address the control of boost converters for the PV and BESS sources and the DAB converter for the thruster, and (iii) propose an uncertainty and disturbance estimator (UDE) concept based on current control algorithms to mitigate MVDC instability due to unpredictable factors and external disruptions. The proposed UDE can actively estimate and compensate for the system disturbance and uncertainty in real time, and thus, both the system tracking performance and robustness can be improved. Simulation studies have been conducted to substantiate the efficacy of the proposed schemes.

MW-Scale High-Voltage Direct-Current Power Conversion for Large-Spacecraft Electric Propulsion

This paper proposes a megawatt (MW)-scale high-voltage (HV) electrical power-conversion element for large-spacecraft electric propulsion (EP) systems. The proposed scheme is intended for long-term and crewed missions, and it is driven by a nuclear electric propulsion (NEP) that acts as a heat source. The scheme includes (i) A two-rotor generator (TRG), (ii) A rectification stage, and (iii) An isolated dual output DC-DC (iDC2) converter. The TRG is a high-reliability electric machine with two rotors, a permanent magnet rotor (PMR), and a wound field rotor (WFR). The PMR has a fixed flux and hence back-EMF, while the back-EMF due to the WFR is controlled by injecting a direct current (DC) into the WFR winding. The total TRG output voltage, which is the sum of voltages due to the PMR and WFR, is controlled over a prescribed region of spacecraft operation. The output of the TRG is rectified and connected to the input of the iDC2 converter. The iDC2 converter uses a three-winding transformer, where the primary winding is fed from the rectified output of TRG, the secondary winding processes the propulsion power to an electric thruster via a high-voltage DC (HVDC) link and a tertiary winding that is connected to the spacecraft’s low-voltage DC (LVDC) power system. Three controllers are proposed for the system: an HVDC voltage controller, an HVDC current controller that controls the voltage and current processed to the thruster, and an LVDC controller that adjusts the current to the LVDC system. Detailed analytical models for the TRG, iDC2 converter, and controllers are developed and verified via simulations under different conditions. The analytical studies are further validated via results from a laboratory prototype.

Cryogenic Liquid Rocket and Nuclear Thermal Rocket: A Survey of the Current Technology and a Comparison for Future Mars Missions

The innovation of reusable rockets holds the potential to revolutionize both Mars scientific exploration and commercial tourism. Thus, this paper aims to indicate the most promising propulsion system for the Mars mission. While SpaceX proposed the cryogenic liquid rocket Starship to achieve Mars colonization by 2050, the viable nuclear thermal propulsion has been outlined by rocketry experts for space travel. This paper surveys the current state-of-the-art of both popular propulsion systems, indicating their advantages and disadvantages. Methods to increase the delivery and cost efficiency are also offered, including on-orbit refueling for cryogenic liquid rockets and various nuclear core variants for nuclear thermal rockets. Subsequently, their comparative analysis is delivered in the discussion, focusing on their transportation efficiency and cost efficiency for future mass missions. With faster travel, greater payload, and reduced propellant consumption, nuclear thermal propulsion has shown to be a brighter prospect and potential than cryogenic liquid rocket. Nevertheless, the author believes a new class of bimodal nuclear thermal electric propulsion system is the most promising option, combining the merits and complementing drawbacks of cryogenic liquid propulsion, nuclear thermal propulsion, and nuclear electric propulsion.

Optimization of Interplanetary Trajectory for Direct Fusion Drive Spacecraft

The Direct Fusion Drive (DFD) technology, which is being developed at present, will allow fast and affordable interplanetary travel. This is a result of the very high specific impulse and the low specific mass of DFD thrusters which outperform more conventional Nuclear Electric Propulsion (NEP), with which it shares the ability of providing a low (albeit higher than in the case of NEP) continuous thrust. It is well known that, to optimize the payload fraction, the thruster should operate in Variable Exhaust Velocity (VEV) mode and that the lower is the specific mass, the higher should be the maximum specific impulse the thruster can produce. A low thrust interplanetary travel, from the orbit around the starting planet to the orbit around the destination planet, can be considered as made of three parts: a first planetocentric phase, a second heliocentric phase and finally a third planetocentric phase; in all of them the trajectory is a sort of a spiral, but while in the first and third the spacecraft makes several (or even a large number) turns about the two planets, the second consists of a fraction of a turn about the Sun. In the first and last one the optimal specific impulse is not much variable and should remain quite low, while in the second one it must go through large variations, reaching a very high value at roughly midway between the planets. To show the potentialities of DFD, three typical fast missions are studied: to the Moon, to Mars and to Titan, showing that this propulsion device will allow humans to reach practically the whole solar system in a reasonable time. Keywords: Interplanetary travel, Human Mars Exploration, Direct Fusion Drive, Trajectory Optimization, Specific Impulse Optimization

Optimization of Interplanetary Trajectory for Direct Fusion Drive Spacecraft

The Direct Fusion Drive (DFD) technology, which is being developed at present, will allow fast and affordable interplanetary travel. This is a result of the very high specific impulse and the low specific mass of DFD thrusters which outperform more conventional Nuclear Electric Propulsion (NEP), with which it shares the ability of providing a low (albeit higher than in the case of NEP) continuous thrust. It is well known that, to optimize the payload fraction, the thruster should operate in Variable Exhaust Velocity (VEV) mode and that the lower is the specific mass, the higher should be the maximum specific impulse the thruster can produce. A low thrust interplanetary travel, from the orbit around the starting planet to the orbit around the destination planet, can be considered as made of three parts: a first planetocentric phase, a second heliocentric phase and finally a third planetocentric phase; in all of them the trajectory is a sort of a spiral, but while in the first and third the spacecraft makes several (or even a large number) turns about the two planets, the second consists of a fraction of a turn about the Sun. In the first and last one the optimal specific impulse is not much variable and should remain quite low, while in the second one it must go through large variations, reaching a very high value at roughly midway between the planets. To show the potentialities of DFD, three typical fast missions are studied: to the Moon, to Mars and to Titan, showing that this propulsion device will allow humans to reach practically the whole solar system in a reasonable time. Keywords: Interplanetary travel, Human Mars Exploration, Direct Fusion Drive, Trajectory Optimization, Specific Impulse Optimization

Advanced Propulsion for Fast Lunar Missions

Half a century ago the Apollo missions carried 12 people to spend a short time on the Moon. After all this time, the space agencies of some countries and a number of private companies committed themselves to restart human lunar exploration, this time with the aim to establish a long-term human presence on our satellite and to start a lunar economy. Most of the planned lunar missions are based on the use of chemical propulsion, which allows to perform the one-way Earth-Moon travel in 3 or 4 days. Recently NASA and DARPA joined forces to develop a nuclear thermal rocket for planetary and lunar missions. In spite of the relatively small distance of the Moon from the Earth, carrying the required payload to the Moon is costly and there is some interest in using more advanced propulsion systems, with the purposes of reducing both the travel time and the cost. In particular, advanced propulsion devices will be developed for travelling to more distant destinations and, when they will be available, they could be used with advantages also for the Moon. The aim of the present paper is to discuss the perspectives which will be opened in the future by the possibility of using nuclear (both thermal and electric) or solar electric propulsion and, in a more distant future, nuclear fusion propulsion to travel to the Moon. A number of examples of lunar missions performed with different types of advanced propulsion show what are the conditions at which these advances can be achieved and which are the constraints that will limit these efforts. Keywords: Lunar Missions, Human Lunar Exploration, Advanced Propulsion, Trajectory Optimization, Specific Impulse Optimization

Near-term strategies to rendezvous with an interstellar object

Interstellar Objects (ISOs) offer a unique opportunity to answer fundamental questions about the nature of objects that originate outside our Solar System. Only two members of this new class of mission target have been identified: asteroid 1I/‘Oumuamua in 2017 and comet 2I/Borisov in 2019. Many more are expected to be found with the introduction of the Vera C. Rubin Observatory motivating the formulation of mission concepts to investigate future ISOs. While fast flyby missions are technically feasible today, rendezvous missions that orbit or even land offer the only means for literal ground truth into the nature of these objects. Mission design and propulsion capability are current challenges due to the brevity of ISO paths through the Solar System. These challenges may be overcome with a trade space of detection capabilities, launch vehicles, trajectory designs, and near-term spacecraft technology development. In particular, the combination of Jupiter gravity assist and nuclear electric propulsion provides the highest likelihood of delivering a spacecraft to match the orbit of an ISO.

Design of a Fast Molten Salt Reactor for Space Nuclear Electric Propulsion

The French National Center for Scientific Research (CNRS) is carrying out design studies on a nuclear electric propulsion (NEP) engine based on a molten salt reactor (MSR). A NEP engine based on liquid nuclear fuel could allow developing a core design with relatively high power densities and temperatures while using simple reactivity control systems and keeping low pressure and temperature gradients in the fuel. Nevertheless, the design work of such an engine poses significant technical challenges and requires the use of advanced numerical simulation tools. Different MSRs for space are currently being studied. In this work, a MSR concept using a fast neutron spectrum is investigated using a multiphysics tool based on a numerical coupling between the OpenFOAM (computational fluid dynamics) and SERPENT 2 (Monte Carlo neutronics) codes. The analysis of this paper is focused on the reactor core coupled neutronic and thermal-hydraulic phenomena. Steady state full-power conditions are calculated for two different fast MSR designs using low-enriched uranium (LEU) and highly enriched uranium. The results show that the proposed core layout and materials allow obtaining a satisfactory temperature distribution in the core (maximal values and gradients) without significant penalization of the reactor operating conditions. A reactivity control strategy excluding the use of control rods is studied for the LEU concept. Transient and safety studies are also performed and show acceptable performance.

Recent concept study for cryogenic fluid management to support opposition class crewed missions to Mars

NASA recently completed a mission concept study to evaluate the feasibility and propulsion technology development requirements for reduced travel duration crewed missions to Mars. A high-level goal of the study was to minimize the health impact on the crew caused by the space environment. This was implemented in the study by limiting the crew to a total of approximately-two years of in-space operations and travel time. For the initial mission, the crew would stay about 30 days on the Martian surface. The propulsive demands of such a mission are immense, and the study identified two advanced propulsion options with the potential to meet the mission requirements—both options rely on nuclear fission to provide efficient propulsive energy. One propulsion option was a nuclear electric propulsion (NEP)/Chem Hybrid, with a reactor and energy conversion system powering xenon propellant ion thrusters to provide an efficient, but lower-thrust, push for most of the mission duration. This concept also relied on a liquid oxygen/liquid methane (LO2/LCH4) chemical propulsion stage to provide high thrust for maneuvers while near the Earth and Mars. The second propulsion option was nuclear thermal propulsion (NTP), in which the reactor heats liquid hydrogen (LH2) propellant to expand through a nozzle for thrust at about twice the efficiency of the best chemical propulsion systems. Both vehicle concepts rely on storing large amounts of cryogenic propellant (either LO2/LCH4 or LH2) for multiple years in space without loss, far exceeding state-of-the-art capability. To enable this new capability, the team assumed the use of several advanced cryogenic fluid management (CFM) technologies and analyzed the integrated system performance. This included considering the vehicle-level effects of the size, mass, and power requirements of these CFM elements. Further, the team evaluated the development required to enable such a mission in the mid-2030 s and determined that it was feasible. The paper elaborates on the assumed CFM technologies, provides key analysis results, and illustrates the feasibility of technology development for the proposed solutions to the CFM challenges for each propulsion concept.

Aerospace Environmental Challenges for Electrical Insulation and Recent Developments for Electrified Aircraft.

The growing trend towards high voltage electrical assets and propulsion in the aeronautics and space industry pose new challenges in electrical insulation materials that cannot be overlooked. Transition to new high voltage electrified systems with unprecedented high levels of voltage, power, and efficiency must be safe and reliable. Improvements in both performance and safety of megawatt power systems is complicated because of the need for additional power transmission wiring and cabling and new safety requirements that have the potential of making the resulting systems heavier. To mitigate this issue, novel lightweight materials and system solutions are required that would result in lower specific weights in the insulator and conductor. Although reduced size and weight of system components can be achieved with new concepts, designs, and technologies, the high voltage (≥300 V) operation presents a significant challenge. This challenge is further complicated when considering the extreme operating environment that is experienced in aircraft, spacecraft, and targeted human exploration destinations. This paper reviews the extreme environmental challenges for aerospace electrical insulation and the needs associated with operating under high voltage and extreme environments. It also examines several recently developed robust lightweight electrical insulation materials that could enhance insulation performance and life. In aerospace, research must consider mass when developing new technologies. The impact of these recent developments provides a pathway which could enable next generation high altitude all electric aircraft, lightweight power transmission cables for a future sustained presence on the Moon and missions to Mars using HV propulsion, such as spacecraft with Nuclear Electric Propulsion systems.

Manned Mars Mission Analysis Using Mission Architecture Matrix Method

With the development of deep space exploration technology, manned missions to Mars are expected to be realized in the near future. However, the journey to Mars requires more supplies and fuel than near-Earth missions, including moon landings. Using the traditional Apollo-style mission method will result in the spacecraft reaching the LEO mass of 2000 tons, which is not easy to achieve. The use of the modular design method and multiple launches will significantly reduce the mass of a single launch. In addition, the use of nuclear power engines (Nuclear Thermal Propulsion, NTP, and Nuclear Electric Propulsion, NEP) can greatly improve propulsion efficiency, reducing the mass of the propellant. This paper uses the Mission Architecture Matrix (MAM) method to concisely and precisely analyze a series of mission architectures in the case of impulse maneuver transfer and low-thrust transfer. The results show that when only chemical propulsion (specific impulse is 440 s) is used, the maximum LEO launch mass in the optimal mission architecture is 325 tons, and the total LEO mass of the system is 1142 tons. With the usage of NTP (specific impulse is 900 s) and NEP (the specific impulse is 6000 s) technology, the maximum LEO launch mass in the optimal mission architecture is only 85 tons, and the total LEO mass of the system is only about 400 tons. Considering the current rocket technology, the total cost is about USD 1149 million US.

Optimum design of nuclear electric propulsion spacecraft for deep space exploration

Future space exploration is now focused on new field-deep space planets. Deep space exploration and the development and utilization of resources are inestimable strategic significance for the country to seize the initiative and command heights of deep space exploration. Nuclear electric propulsion (NEP) systems convert heat from the fission reactor to electrical power and then use the electrical power to produce thrust. Compared with traditional propulsion technology, the NEP system with variable Isp can be used for several applications. The NEP spacecraft is more suitable for deep space exploration missions due to the advantages of high specific impulse, high power, and long life. This paper analyzes the relationship between the transfer travel time, specific mass, power, and payload ratio of the NEP spacecraft through simple performance models. Use the mass optimization and specific mass optimization models based on the NEP system composition and small thrust orbit theory to maximize the payload ratio for a given transfer time and technological characteristics. Finally, applied this NEP model to discuss the feasibility of Mars, Jupiter, and Saturn transfer missions and compared it with the Tianwen-1, Juno, and Cassini–Huygens spacecraft, respectively. Results show that when the NEP spacecraft specific mass reaches 4.77 kg/kW, the Earth–Mars transfer time can be changed to 331.31 days, the payload increases to 1650 kg, the transfer time for Jupiter and Saturn mission would be shorted to 661.31 days and 1131.31 days, the corresponding payload significantly increases which achieved to 1270 kg and 2981 kg. The nuclear electric propulsion spacecraft dramatically improves the detection capability of the spacecraft and provides a reference for the feasibility demonstration and subsequent design of the deep space and the extrasolar planet exploration.

Trajectory Analysis and Optimization of Hesperides Mission

A challenging problem from a technological viewpoint is to send a spacecraft at a distance of about 600 au from the Sun, comparable with that of the Sun’s gravitational focus (that is, the general relativistic focusing of light rays, whose minimum solar distance is obtained when the light rays are assumed to graze the Sun’s surface), and reach it in a time interval on the order of a human working lifetime. A suitably oriented telescope at that distance would be theoretically able to observe exoplanets tens of light years far away and possibly to discover new life forms. The transfer trajectory of this mission is rather complex and requires a close selection of a suitable propulsion system, which must be able to provide the probe with the necessary energy to cruise at a velocity greater than 10 au/year. An effective outline of the these concepts is given by the Hesperides mission, originally proposed by Matloff in 2014. An interesting aspect of this mission proposal is the combination of a nuclear electric propulsion system and a classical solar sail that are jointly exploited to reach the necessary solar system escape velocity. However, the trajectory analysis reported by Matloff is very simplified and is essentially concentrated on a rough estimate of the time required by the spacecraft to reach a distance of 600au. Starting from the Hesperides baseline mission proposal, including the vehicle mass distribution, the aim of this work is to give a detailed mission analysis in an optimal framework. In particular, the spacecraft minimum time trajectory is calculated with indirect methods and a parametric analysis is made to highlight the impact of the main design parameters on the total flight time. The simulations show a substantial reduction of the mission time when compared with the original study.

Nuclear Power Concepts and Development Strategies for High-Power Electric Propulsion Missions to Mars

Under the Mars Transportation Assessment Study, NASA and the U.S. Department of Energy are performing analyses and generating concepts for crewed nuclear electric propulsion (NEP) missions to Mars. This paper presents the results of trade studies and concept development for the nuclear electric power system, consisting of the fission reactor, radiation shielding, power conversion, heat rejection, and power management and distribution (PMAD). The nuclear power team completed trade studies to evaluate different reactor and power conversion technologies and developed preliminary concepts for the crew shielding, waste heat radiators, and PMAD. The initial results suggest that a modified terrestrial microreactor combined with supercritical CO2 Brayton conversion could be used to perform the crew and cargo missions with satisfactory performance and modest risk. The paper includes preliminary development strategies that could bring the NEP technology to fruition for Mars missions in the late 2030s or early 2040s.

Enabling a New Generation of Outer Solar System Missions: Engineering Design Studies for Nuclear Electric Propulsion

Enabling a New Generation of Outer Solar System Missions: Engineering Design Studies for Nuclear Electric Propulsion

Novel Deep Space Nuclear Electric Propulsion Spacecraft

The space industry is expanding at an increasing rate. While most efforts are currently focused on Earth and lunar orbits, it is only a matter of time before affordable exploration missions into deep space become more prevalent. Nuclear electric propulsion (NEP) with large quantities of power have been theorized for such missions with many advantages over traditional solar panels and radioisotope power sources. Key among NEP issues has been the power conversion system, often falling upon dynamic cycles over solid-state options like thermoelectric generators (TEGs) because of low efficiencies. Howe Industries has conceptualized a deep space probe capable of transporting cube satellites (CubeSats) and other payloads to deep space utilizing NEP based on an advanced TEG power conversion system with efficiencies that would challenge traditional dynamic power conversion cycles. Experimentation at a TRIGA research reactor has shown a potential for 20 to 50 times increase in electrical conductivity of potential thermoelectric materials, which would correlate to large increases in efficiencies over traditional TEGs.

Electric Propulsion Methods for Small Satellites: A Review

Over 2500 active satellites are in orbit as of October 2020, with an increase of ~1000 smallsats in the past two years. Since 2012, over 1700 smallsats have been launched into orbit. It is projected that by 2025, there will be 1000 smallsats launched per year. Currently, these satellites do not have sufficient delta v capabilities for missions beyond Earth orbit. They are confined to their pre-selected orbit and in most cases, they cannot avoid collisions. Propulsion systems on smallsats provide orbital manoeuvring, station keeping, collision avoidance and safer de-orbit strategies. In return, this enables longer duration, higher functionality missions beyond Earth orbit. This article has reviewed electrostatic, electrothermal and electromagnetic propulsion methods based on state of the art research and the current knowledge base. Performance metrics by which these space propulsion systems can be evaluated are presented. The article outlines some of the existing limitations and shortcomings of current electric propulsion thruster systems and technologies. Moreover, the discussion contributes to the discourse by identifying potential research avenues to improve and advance electric propulsion systems for smallsats. The article has placed emphasis on space propulsion systems that are electric and enable interplanetary missions, while alternative approaches to propulsion have also received attention in the text, including light sails and nuclear electric propulsion amongst others.

A review of high-power magnetoplasmodynamic electric propulsion designs and studies of RSC Energia

At the initiative of S.P.Korolev, in 1959, Special Design Bureau No.1 (now RSC Energia) established the High-temperature Power Engineering and Electric Propulsion Center which was tasked with development of nuclear electric propulsion for heavy interplanetary vehicles. Selected as the source of electric power was a nuclear power unit based on a thermionic converter reactor, and selected as the engine was a stationary low-voltage magnetoplasmodynamic (MPD) high-power (0.5–1.0 MW) thruster which had thousands of hours of service life. The paper presents the results of extensive efforts in research, development, design, materials science experiments, and tests on the MPD-thruster, including the results of development and 500-hours life tests of an MPD-thruster with a 500-600 kW electric power input that used lithium propellant. The world’s first lithium 17 kW MPD-thruster was built and successfully tested in space. The paper points out that to this day nobody has surpassed the then achievements of RSC Energia neither in thruster output during long steady-state operation, nor in performance and service life. Key words: Martian expeditionary vehicle, nuclear electric rocket propulsion system, electric rocket thruster, magnetoplasmodynamic thruster, lithium, cathode, anode, barium, electric propulsion tests in space.