Nuclear Electric Propulsion System Research Articles

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.

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.

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.

Trajectory Optimization of a Low-Thrust Geostationary Orbit Insertion for Total Ionizing Dose Decrease

The paper considers a method of the dose exposure decrease from the charged particles of Earth’s radiation belts (ERBs) affecting a reusable orbital transfer vehicle, which is launched from a low circular orbit into geostationary orbit using a nuclear electric propulsion system. The main idea of the method consists in numerical continuation of a solution of the minimum time problem with respect to an ionizing radiation dose accumulated at the end of a transfer. To do this, equations of motion of the orbital transfer vehicle are supplemented by an additional equation for the radiation dose, and the boundary condition for the dose at the right end is introduced. In calculating the dose, the AE8/AP8 MIN, AE8/AP8 MAX, and AE9/AP9 models of fluxes of charged particles of ERBs were used. By changing the insertion trajectory shape, it became possible to lower the radiation dose by 25–38% relative to the minimum time trajectory. At the same time, the transfer time increased no more than by 7% of the minimum time of launching into geostationary orbit, and the characteristic velocity expenses increased by 320–560 m/s.

Solar electric propulsion for human mars missions

While solar electric propulsion (SEP) is being widely considered for cargo transport to Mars, its value for propelling fast human missions is often viewed as marginal. This conclusion is driven by the high electric power requirement (multi megawatts) of a fast human spacecraft, coupled to the low power density of traditional solar arrays. For these applications, nuclear electric propulsion (NEP) appears to provide a better electric alternative. However, recent progress in the field of thin film photovoltaic cells and large deployable structures may, at least in the short-term, challenge this conclusion. Although ultimately NEP systems might very well become the mainstay of fast human deep space transport, we examine the human SEP option as an attractive intermediate path on this journey, one that capitalizes on the rapid evolution of the solar array technology being experienced today. We investigated the challenges of building suitably large, lightweight, solar arrays to produce the required electric power for both cargo and human interplanetary spacecraft and examined the advantages of such SEP architectures in the context of a long-stay human Mars mission. Within this framework, we present some conclusions regarding the prospects for this technology.

Analysis of the Requested Perfection of a Nuclear Electric Propulsion System for a Mars Mission with a 2-Year Duration

Abstract—The results of the design and ballistic analysis of a 2-year manned mission to Mars are presented. Dependences of the maximum permissible specific mass of the electric power and propulsion system on the spacecraft (SC) mass in the initial earth orbit are obtained. It is shown that, for the analyzed set of characteristics of SC systems with the SC mass in the initial orbit of 200 t, to implement a Mars mission with a 2-year duration, the required perfection of the electric power and propulsion system should be ideally high (the specific mass of the electric power and propulsion system should be no more than 3.26 kg/kW). Increasing the mass in the initial orbit leads to a relaxation of the requirements for the perfection of a transport system. If the mass in the initial orbit is increased to 475 t, then the maximum possible specific mass of the electric power and propulsion system grows to 11 kg/kW. The optimal electric power of the nuclear power plant with a change in the initial mass within the mentioned range (from 200 to 475 t) increases from 7.7 to 11.7 MW. The optimal specific impulse of the electric propulsion falls from 9000 s (this value is accepted as the maximum permissible) to 6880 s. It is shown that, if the specific mass of the electric power and propulsion system is 5, 7.5, and 10 kg/kW, then, to implement the mission, the SC mass in the initial orbit should be no less than 234.1, 305.4, and 415.5 t, respectively.

Neutronics and Thermal Hydraulics Analysis of a Conceptual Ultra-High Temperature MHD Cermet Fuel Core for Nuclear Electric Propulsion

Nuclear electric propulsion (NEP) offers unique advantages for the interplanetary exploration. The extremely high conversion efficiency of magnetohydrodynamics (MHD) conversion nuclear reactor makes it a highly potential space power source in the future, especially for NEP systems. Research on ultra-high temperature reactor suitable for MHD power conversion is performed in this paper. Cermet is chosen as the reactor fuel after a detailed comparison with the (U,Zr)C graphite-based fuel and mixed carbide fuel. A reactor design is carried out as well as the analysis of the reactor physics and thermal-hydraulics. The specific design involves fuel element, reactor core and radiation shield. Two coolant channel configurations of fuel elements are considered and both of them can meet the demands. The 91 channel configuration is chosen due to its greater heat transfer performance. Besides, preliminary calculation of nuclear criticality safety during launch crash accident is also presented. The calculation results show that the current design can meet the safety requirements well.

Shadow Radiation Shield Required Thickness Estimation for Space Nuclear Power Units

The paper concerns theoretical possibility of visiting orbital transport vehicles based on nuclear power unit and electric propulsion system on the Earth's orbit by astronauts to maintain work with payload from the perspective of radiation safety.There has been done estimation of possible time of the crew's staying in the area of payload of orbital transport vehicles for different reactor powers, which is a consistent part of nuclear power unit.

One Variant of Manned Mission to Mars with a Nuclear Electric Propulsion

In this paper, one possible way for implementing a manned mission to Mars is examined. Typical peculiarities of the mission are as follows: the nuclear electric propulsion; relatively low mass of the spacecraft at a low Earth orbit (200 tons) and the crew time in flight is high (900-1000 days). Space mission analysis of the chosen variant is performed. As an optimization criterion, the authors chose the fuel mass required for the flight. Under examined problem definition such mass minimization is equivalent to maximal final mass of the spacecraft and maximal permissible total mass of power and electric propulsion systems. The authors show that to implement the examined manned mission, it is necessary to create the nuclear electric power and electric propulsion systems with a specific mass lower than 12.5 kg/kW under propulsion efficiency of 0.8, specific mass of the system for propellant storage of 0.05 and manned spacecraft complex mass of 52.1 tons. Under propulsion efficiency of 0.7, specific mass of power-propulsion should be lower than 10.9 kg/kW.

Mars Surface Exploration Caching from an Orbiting Logistics Depot

E XPLORATION missions on Earth have used the principle of caching for centuries. Caching involves prepositioning discrete quantities of supplies at specific locations along a route in order to lighten the burden or extend the range of an exploring party. Historical missions that made use of caching on Earth are the Lewis and Clark expedition in 1803–1806 [1] and the ill-fated Imperial Trans-Antarctic Expedition of 1914–1917 [2]. Caching may play an important role in the future of manned planetary exploration as well. Without it, exploration sorties are inherently limited to a specific distance from a base or landing location given by the type of transportation vehicles used, the size of the crew, the mix and quantity of consumables, operational rules, and the nature of the terrain. For manned lunar missions, for example, this distance has been estimated at 3 km for exploration on foot, 5–15 km with unpressurized rovers, and up to 30 kmwith pressurized rovers [3]. In part, exploration range is determined by the quantity of consumables that can be carried along, limiting the amount of useful exploration that can be done at distances greater than 30 km from the landing site or base. Similar issues exist for manned Mars exploration. This Note explores the concept of storing caches in a logistics depot on orbit and deliberately deploying them to the surface to extend human exploration range.Wewill useMars as the motivating context, but the idea may apply to other situations as well. We first introduce the concept and then perform initial sizing of such a depot. This Note does not explore the full tradespace of orbital depots and caching architectures. Caching in space exploration can be implemented as two different architectures: direct entry from inspace trajectory and orbital entry from a low Mars orbit. While the direct entry does not contain the Martian orbit insertion maneuver and can be more efficient in terms of required V, this Note only considers the orbital entry because of the following two reasons. First, the short time gap between deployment decision and actual landing of the supply unit provides meaningful flexibility to operate on orbit. Secondly, advanced propulsion systems such as a nuclear electric propulsion system (NEP) and solar electric propulsion can significantly reduce the cost resulting from the orbit insertion, and the efficiency gap gets very small. Let us consider an orbiting logistics depot that is composed of a cluster of supply units deployed to aMartian orbit. Each prepackaged supply unit contains supply items (fuel and consumables) that allow extending the range of humans exploring the surface of Mars. Figure 1 shows the orbiting depot concept and its concept of operations. Typically, one or more surface vehicles start exploration from a base and return to the base before all fuel and consumables (e.g., fuel, food, water, and oxygen) have been expended. This is the case of an unassisted surface route, yielding a total distanceD. In this case, the amount of fuel and consumables that can be used during a single route is determined by the capacity of the surface exploration vehicle. Now suppose that there is a logistics depot in Martian orbit and it is possible to command the depot to drop a supply unit (pod) filled with a known mix and quantity of consumables corresponding approximately to a vehicle’s capacity. If the supply unit lands at such a location that the surface exploration vehicle can reach the unit before it runs out of consumables, the vehicle’s range can effectively be doubled to 2D. This is the case of the depot-assisted surface route shown in Fig. 1 (right) along with a simulated pod landing error ellipse. In the following sections we discuss the most important design decisions for such an orbiting depot, including orbit selection, individual supply unit design, and depot operations. These calculations are meant to establish a realistic “strawman” for the concept. Studies regarding detailed design in the context of specific surface exploration campaign strategies are left for future work.

Analysis of Electric Propulsion System for Exploration of Saturn

Exploration of the outer planets has experienced new interest with the launch of the Cassini and the New Horizons Missions. At the present time, new technologies are under study for the better use of electric propulsion system in deep space missions. In the present paper, the method of the transporting trajectory is used to study this problem. This approximated method for the flight optimization with power‐limited low thrust is based on the linearization of the motion of a spacecraft near a keplerian orbit that is close to the transfer trajectory. With the goal of maximizing the mass to be delivered in Saturn, several transfers were studied using nuclear, radioisotopic and solar electric propulsion systems.

Proposal of Space Reactor for Nuclear Electric Propulsion System

Currently, the solar battery, the chemical cell, and the RI battery are used for the energy source in space. However, it is difficult for them to satisfy requirements for deep space explorations. Therefore, other electric power sources which can stably produce high electric energy output, regardless of distance from the sun, are necessary to execute such missions. Then, we here propose small nuclear reactors as power sources for deep space exploration, and consider a conceptual design of a small nuclear reactor for Nuclear Electric Propulsion System. It is found from nuclear analyses that the Gas-Cooled reactor could not meet the design requirement imposed on the core mass. On the other hand, a light water reactor is found to be a promising alternative to the Gas-Cooled reactor.

Pellet bed reactor concepts for nuclear propulsion applications

Pellet bed reactor (PeBR) concepts have been developed for nuclear thermal and nuclear electric propulsion, and bimodal applications. This annular core, fast spectrum reactor offers many desirable design and safety features. These features include high-power density, small reactor size, full retention of fission products, passive decay heat removal, redundancy in reactor control, negative temperature reactivity feedback, ground testing of the fully assembled reactor using electric heating and nonnuclear fuel elements, and the option of fueling on the launch pad or fueling and refueling in orbit. In addition to these features, the concepts for nuclear electric propulsion and for bimodal power and thermal propulsion have no single point failure. The average power density in the reactor for nuclear thermal propulsion ranges from 2.2 to 3.3 MW/I and for a 15-MWe nuclear electric propulsion system the total power system specific mass is about 3.3 kg/kWe. The bimodal-PeBR system concepts offer specific impulse in excess of 650 s, tens of Newtons of thrust, and total system specific power ranging from 11 to 21.9 We/kg at the 10- and 40-kWe levels, respectively. 35 refs.

Nuclear propulsion for space exploration

Nuclear propulsion has been identified as a key technology by several groups that have studied how to implement the Space Exploration Initiative to go back to the Moon, establish human presences on Mars and explore space beyond Mars. This paper will describe the results of some of the recent studies of the application of both nuclear electric and nuclear thermal propulsion systems in space exploration. This paper will also identify some of the issues that require further study and that have a significant effect on the propulsion system design and selection.

Comparison of Advanced Propulsion Capabilities for Future Planetary Missions

This paper summarizes unmanned planetary performance (payload and trip time) of Shuttle-based advanced propulsion systems for 1980-90 missions analyzed as part of the recent NASA/AEC Advanced Propulsion Comparisons Studies. Propulsion system designs and condensed results from over 300 propulsion/mission combinations are discussed. Chemical rocket (CRP), solar electric (SEP), nuclear rocket (NRP), and nuclear electric (NEP) propulsion systems are all considered. In terms of missions flown, total flight time, and number of Shuttle launches required, NEP provides the best performance. Relative to NEP, it is shown that NRP, SEP, and CRP degrade mission performance by 20%, 40%, and 50%, respectively, at nominal payloads.

Thermodynamic characteristics of brayton cycles for space power

for planetary landing missions, TR 32-231, Jet Propulsion Lab., Pasadena, Calif. (May 1, 1962). 3 Speiser, E. W., Performance of nuclear-electric propulsion systems in space exploration, TR 32-159, Jet Propulsion Lab., Pasadena, Calif. (December 15, 1961). 4 Stearns, J. W., Electric propulsion requirements for planetary and interplanetary spacecraft, TR 32-403, Jet Propulsion Lab., Pasadena, Calif, (to be published). 5 Hayes, R. J., Hoffman, M. A., and Stuhlinger, E., The manned electric rocket challenge, Preprint 2222-61, ARS Space Flight Report to the Nation, New York (October 1961). 6 Melbourne, W. G. and Sauer, C. G., Optimum thrust programs for power-limited propulsion systems, TR 32-118, Jet Propulsion Lab., Pasadena, Calif. (June 15, 1961). 7 Volkoff, J. J., Protection requirements for the resistance of meteorite penetration of interplanetary spacecraft systems, TR 32-410, Jet Propulsion Lab., Pasadena, Calif, (to be published). 8 Davis, J. P., A nuclear reactor concept for electric propulsion application, TR 32-385, Jet Propulsion Lab., Pasadena, Calif. (March 4, 1963).