Electric space propulsion systems

Most of marine electric propulsion ships are propelled by large electric motors. The speed control of the propulsion motor on the electric propulsion vessel is operated by adjusting the frequency. The power system has the inverter installed semiconductor element, such as IGBT or IGCT, for high-speed switching. However, the inverter for power conversion unit of the electric power propulsion on large vessels has fundamentally thermal problems of installed semiconductor, such problems could influence to considerable loss by switching and conducting the inverter in the power system transferring large current with high voltage. In order to calculate power loss and thermal loss of the inverter, this paper sets actual parameters of the electric propulsion system of the large ship in the simulation program considering condition of ship operation. The simulation proves that it is possible to analyze both power loss and thermal loss in the diodes and transistor during switching of the inverter in comparison with field oriented control and direct torque control. As a result, large power loss and thermal loss of switching and conducting the inverter by field oriented control considerably occurs and can also impact damages toward the power electronic devices by the thermal loss generated.

The paper investigates the utility of how nuclear thermal rocket (NTR) upper stage propulsion could substantially reduce the interplanetary transit time while greatly enhancing the economics of performing the manned Mars mission. au ECCV GTO IMLEO Isp kg kg/kw km km/sec kw LEO MEV mw NTR sec !N Nomenclature astronomical unit Earth capture crew vehicle geosynchronous transfer orbit initial mass in low Earth orbit specific impulse kilogram kilograms per kilowatt kilometer kilometers per second kilowatt low Earth orbit Mars excursion vehicle megawatt nuclear thermal rocket seconds delta V 1. Background Initial responses to the launch cost issue have been to suppose the development of heavy lift launch vehicles that are assumed to have fundamental launch cost savings. In general, most studies have isolated launch vehicle systems from the interplanetary vehicle assembled in low Earth orbit. This paper investigates a potential synergism between applying nuclear thermal rockets as both upper stages of the launch vehicle and as departure propulsion for a Mars vehicle departure stage. II. InterPlanetary Vehicle Description The mission to Mars is a propulsion technology dominated mission. The velocity change requirements (llV or delta V) are very large and very sensitive to trip time constraints. Due to this, there has been a heavy emphasis on high specific impulse (Isp) propulsion systems for performing the manned missions to Mars. The fundamental problem of high Isp systems is that they are generally thermally limited by material temperature limits. This issue is eliminated by electric propulsion systems that use electro-magnetic forces for accelerating a working fluid, but are thrust limited due to practical limitations in power supply mass. More exotic propulsion systems tend to use plasma working fluids that can only be contained by such methods as magnetic nozzles. The Space Exploration Initiative (SET) has a long range goal to perform a manned mission to Mars. Preliminary design goals of an interplanetary vehicle for this mission stress extremely high propulsion requirements to minimize transit times while allowing significant stay times on Mars. Related to this key trade between propulsion performance and trip time is the space transportation economics of 20 or more Earth launches required of the current launch vehicle architecture to place adequate vehicle mass in low Earth orbit to perform a single manned mission to Mars. The large initial mass in low Earth orbit (IMLEO) to perform the manned Mars mission is recognized as a dominant mission cost driver using the current launch infrastructure. IMLEO has been used as a preliminary figure of merit to evaluate various design approaches. * Senior Advanced Propulsion Engineer, Member AIAA Most analysis to date has focused on nuclear thermal rockets (Isp ... 1000 sec) and megawatt class electric propulsion systems (Isp ... 5000+ sec). The vehicle strategies have been to split the mission between a one way only cargo mission to Mars and a fast transfer manned vehicle. Mars rendezvous is essential for mission success, and techniques to slow at Earth return have used Earth capture crew vehicles (ECCV) with very high excess velocities (9+ km/sec) that are ballistically braked. Analyzed here is a very simple single vehicle. It uses an NTR driven, high thrust, Earth departure stage that is dropped once the vehicle has entered a fast Mars transfer ellipse. Once the vehicle has exited the deep Earth gravity well, it relies solely on a 5000 second Isp electric propulsion system for capture at Mars, departure from Mars, and final mission capture at Mars with zero excess velocity (a conservative assumption) The payload to Mars is broken into two elements. The ECCV is assumed to weigh 40,000 kg for a crew of four, and is the crew's interplanetary travel habitation module. The Mars excursion vehicle (MEV) is the element used to drop the crew from Mars orbit to the planet surface, be the crew's planetary habitation module, carry all science equipment, and carry the ascent module for returning to Mars orbit. The MEV has a baseline weight of 84,000 kg for a nominal 30 day planetary stay time and is scaled with the squareroot of normalized stay time. The MEV is not returned to Earth. Consumables are estimated at 10 kg per person per mission day. optimized for minimum l1V by adjusting the up leg perihelion and aphelion while maintaining the constraints of Earth rendezvous and travel time. Figure 2 (with the radius in millions of kilometers) shows a transfer ellipse for the up leg for the 320 day travel time, 80 day Mars stay time, 0 degree mean Mars true anomaly case. Figure 3 shows the down leg of that mission. The mission l1V requirement of 29.1 krn/sec is divided into 4.2 krn/sec for Earth departure (in heliocentric space), 6.0 krn/sec for Mars capture, 7.4 krn/sec for Mars departure, and 11.5 krn/sec for Earth capture. Transfer Ellipse Up. III. Interplanetary Mission Modeling The problem of analyzing the interplanetary mission can be driven by very exacting issues such as actual planetary configurations for particular mission opportunities, the requirements for travel time, and the Mars stay time. The issue that needs near term resolution, however, is the general efficacy of various propulsion and/or vehicle configuration options for the Mars mission in general. To answer these type of questions, it was necessary to generalize the Mars transfer problem. To do so, simplifying assumptions were that the Mars and Earth orbits were coplanar, and that the Earth orbit was circular. The nominal planetary configurations for Mars, Earth, and Venus are shown in Figure 1. The radius units are in astronomical units (Earth radius =1 au) and the angular units are the true anomaly of the orbital positions in degrees (Mars perihelion = 0 degrees). tojMJoC:: ~ •• ~: ••• . -

Solar sails have long been known to be an attractive method of propulsion in the inner solar system if the areal density of the overall spacecraft (S/C) could be reduced to ∼10 g/m2. It has also long been recognized that the figure (precise shape) of useful solar sails needs to be reasonably good, so that the reflected light goes mostly in the desired direction. If one could make large reflective surfaces with reasonable figure at an areal density of ∼10 g/m2, then several other attractive options emerge. One is to use such sails as solar concentrators for solar-electric propulsion. Current flight solar arrays have a specific output of ∼100W/kg at 1 Astronomical Unit (AU) from the sun, and near-term advances promise to significantly increase this figure. A S/C with an areal density of 10 g/m2 could accelerate up to 29 km/s per year as a solar sail at 1 AU. Using the same sail as a concentrator at 30 AU, the same spacecraft could have up to ∼45 W of electric power per kg of total S/C mass available for electric propulsion (EP). With an EP system that is 50% power-efficient, exhausting 10% of the initial S/C mass per year as propellant, the exhaust velocity is ∼119 km/s and the acceleration is ∼12 km/s per year. This hybrid thus opens attractive options for missions to the outer solar system, including sample-return missions. If solar-thermal propulsion were perfected, it would offer an attractive intermediate between solar sailing in the inner solar system and solar electric propulsion for the outer solar system. In the example above, both the solar sail and solar electric systems don't have a specific impulse that is near-optimal for the mission. Solar thermal propulsion, with an exhaust velocity of the order of 10 km/s, is better matched to many solar system exploration missions. This paper derives the basic relationships between these three propulsion options and gives examples of missions that might be enabled by such hybrids.

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.

NASA has completed a preliminary mission and systems study of nuclear electric propulsion (NEP) systems for “split‐sprint” human exploration and related robotic cargo missions to Mars. This paper describes the study, the mission architecture selected, the NEP system and technology development needs, proposed development schedules, and estimated development costs. Since current administration policy makers have delayed funding for key technology development activities that could make Mars exploration missions a reality in the near future, NASA will have time to evaluate various alternate mission options, and it appears prudent to ensure that Mars mission plans focus on astronaut and mission safety, while reducing costs to acceptable levels. The split‐sprint nuclear electric propulsion system offers trip times comparable to nuclear thermal propulsion (NTP) systems, while providing mission abort opportunities that are not possible with “reference” mission architectures. Thus, NEP systems offer short transit times for the astronauts, reducing the exposure of the crew to intergalactic cosmic radiation. The high specific impulse of the NEP system, which leads to very low propellant requirements, results in significantly lower “initial mass in low earth orbit” (IMLEO). Launch vehicle packaging studies show that the NEP system can be launched, assembled, and deployed, with about one less 240‐metric‐ton heavy lift launch vehicle (HLLV) per mission opportunity ‐ a very large cost savings! Technology development cost of the nuclear reactor for an NEP system would be shared with the proposed nuclear surface power systems, since nuclear systems will be required to provide substantial electrical power on the surface of Mars. The NEP development project plan proposed includes evolutionary technology development for nuclear electric propulsion systems that expands upon SP‐100 (Space Power ‐ 100 kw) technology that has been developed for lunar and Mars surface nuclear power, and small NEP systems for interplanetary probes. System upgrades are expected to evolve that will result in even shorter trip times, improved payload capabilities, and enhanced safety and reliability. Non‐nuclear technology development for the NEP system is estimated to cost about $721 M (1993 $). Nuclear technology development costs are not included in the costs in this report, since these costs will be incurred in the nuclear surface power development program. NEP Phase A/B studies are estimated to cost about $154 M. Flight system hardware development (Phase C/D) is estimated to cost about $2.8 B, and fabrication of flight hardware is estimated to be about $7.8 B for four mission opportunities in 2009, 2011, 2014, and 2016.

Multi-revolution transfer for heliocentric missions with solar electric propulsion

ABSTRACTThe increasing environmental requirements in the air transport sector pose great challenges to the aviation industry and are key drivers for innovation. Besides various approaches for increasing the efficiency of conventional gas turbine engines, electric propulsion systems have moved into the focus of aviation research. The first electric concepts are already in service in general aviation. This study analyses the potentials of electric and turbo hybrid propulsion systems for commercial aviation. Its purpose is to compare various architectures of electrical powertrains with a conventional turboprop on a regional aircraft, similar to the ATR 72, on engine and flight mission levels. The considered architectures include a turbo-electric (power controlled and direct driven), hybrid-electric (serial and parallel) and a pure electric concept. Their system weights are determined using today’s technology assumptions. With the help of performance models and flight mission calculations the impact on fuel consumption, CO ${}_{2}$ emissions and aircraft performance is evaluated.

Abstract: In this paper, describes about advancement in solar-based power system in satellite. A thrust which provides a back force to satellites that move in upward direction and get satellites help to generate power in space. This power transmits to the earth via wireless power transmission and such type of satellite called space based solar power satellite. In recent time, electrical propulsion system has used mostly in place of chemical propulsion system for satellite. Solar energy collected from the space and converts into DC then transmit power from the space to earth wireless and then send to electric grid system. Solar energy is also used by the thruster elements in DC. Keywords: WPT (wireless power transmission), SBSP, propellant, propulsion Cite this Article Divyanshi Srivastava, S.K. Srivastava. Solar-based Satellite Power System: An Overview. Journal of Thermal Engineering and Applications. 2019; 6(1): 20–23p.

Combustion engines are increasingly being regarded as unsustainable in the long-term, because of their negative impact on the environment (e.g. pollution, green-house gas production, and global warming). This has generated worldwide interest in propulsion systems based on renewable alternative energy sources for the future. Fuel cell technology is a promising alternative power source because of their high specific energy, efficiency, and reliability. Hydrogen proton exchange membrane fuel cell (PEMFC) in particular produces zero carbon emissions by having only water vapor as the exhaust. Although there has been much research by automotive industries in developing fuel cell hybrid electric vehicles (FCHEV), fuel cell research for aircraft application is relatively new. Therefore, there is a pressing need for research related to development of aircraft fuel cell electric propulsion systems. Universiti Teknologi MARA (UiTM) is conducting static experiments on different configurations of fuel cell electric propulsion systems. The objective of this study is to understand the behavior of a PEMFC propulsion system under a ground-based static test. A 1 kW PEMFC was used as the main power source for a brushless DC motor electric propulsion system. The electrical characteristics, rotational speed, and thrust data were presented for two different electrical propellers. Analyses of the results were used to characterize the effectiveness of the fuel cell system and its balance of plant. The results were beneficial as a predictive method on defining the optimum electric propulsion system performance needed for future actual flight development.

A brief status review of the advanced launch system (ALS) and its associated advanced development program are presented to demonstrate maturation of those technologies that will help meet the overall operability and cost goals. The electric power and actuation systems are highlighted as a specific technology ready not only to meet the stringent ALS goals (cryogenic fuel valves and thrust vector controls with peak power demands to 75 hp), but also those of other launch vehicles, military and civilian aircraft, lunar/martian vehicles, and a multitude of commercial applications. >

Electric propulsion systems are nowadays one of the most important ways to overcome the modern energy paradigm. Among other systems, electric boats can take advantage of using electric power increasing the efficiency of the propulsion chain and reducing the pollutant gas emissions. This paper presents and discusses global aspects of a boat propulsion system investigated using a low scale system. The proposed chain is modeled and analyzed using two different battery types (lead acid and Ni-MH) and a power converter - DC motor set controlled according with the specifications of such systems. Using classical approaches (as for instance, for the DC motor model), the proposed solution is analyzed implementing a speed control system with a sliding mode controller.

Application of electrical propulsion for an active debris removal system: a system engineering approach

20 - Electric propulsion systems

Fast sample return from the outer Solar System would open an entirely new avenue for space science, but the vast distances make this a daunting task. The achievable transit velocity and the need for extra propellant on the return trip limit the feasibility of returning extraterrestrial samples to Earth. To keep the mission duration short enough to be of interest, sample return from objects farther out in the Solar System requires increasingly higher velocities. High specific impulse, electric propulsion reduces the propellant required for the outbound and return trips, but decelerating the spacecraft at the inner Solar System from high velocity still involves a long, inward spiral trajectory. The use of solar sails to rapidly decelerate incoming sample capsules and eliminate propellant is explored in this paper. The sail is essentially a ''solar parachute'' used for braking at the end of the interplanetary return flight, permitting a higher transit speed and truncating the deceleration spiral. In this application the sail is relatively small and manageable since only the sample capsule and its sail are decelerated. A comparison is made between using all-electric propulsion versus combining electric propulsive acceleration with sail deceleration for sample return from the distances of Saturn, Uranus, and Pluto. Solar-sail braking dramatically reduces the return flight time by one-third or more compared to using electric rocket deceleration. To elucidate the technology requirements, wide ranges for both the loaded sail density and electric propulsion specific mass are considered in this initial parametric study.

The paper gives examples of orbital maneuver execution of spacecraft (SC) using orbit correction propulsion system (OCPS) based on electric propulsion (EP). The high velocity of the propellant flow achieved in EP allows for orbital maneuvers with significantly lower propellant flow rate than in conventional propulsion systems. The time required to perform the orbital maneuvers using EP is connected with the available on-board power. It is typically much greater than such time needed for conventional jet propulsions. The advantages of EP are realized if mission allows a long-time operation of the propulsion system. Since the early 1970s stationary plasma thrusters (SPTs) developed on the base on the concept proposed by A.I. Morozov are used on the SC Meteor. With the help of OCPS based on SPT EOL-1, SC Meteor was installed on the conventionally synchronous orbit, which provides a fixed grid of tracks with a period of T = 102.31 min. In this case, a complete overview of all daily Earth’s surface with the bandwidth of 2900 km is obtained. In recent decades, there has been a steady trend toward miniaturization of space equipment, which requires the development of acceptable thrusters to meet new requirements. Typical total impulses of thrust required for OCPS are reduced several times. At present, a number of space constellations are being developed in Russia based on small SC with a mass from 60 to 500 kg. VNIIEM Corporation creates a space constellation IONOZOND, designed to monitor the geophysical conditions. The description of the IONOZOND constellation is given and the options for using of various OCPS, in particular, based on SPT, ion and pulsed plasma thrusters are considered. It is shown that their use in small SC can significantly increase the economic efficiency of remote sensing orbital constellations.