Using Simscape™ to Support System-Level Design

Full electric propulsion system becomes popular recently and getting more common in the communication satellite industry. Full electric or hybrid propulsion selection causes results in significant reduction in satellite propellant mass. Conventional full chemical satellite propellant mass is roughly 2/3 of launch mass so this huge reduction in propellant mass changes satellite design and preferences. Depending on satellite size and launch vehicle performance, full electric satellite requires 3-4 times less propellant compared to full chemical satellite and 1-2 times less propellant compared to the hybrid system. Satellite launch mass can be reduced by 40% for full electric propulsion and 15% for hybrid propulsion hence launch cost can be decreased by 40% and 15 % with the benefit of reducing propellant mass. It is possible to extend satellite communication capacity average 30 transponders for full electric and average 10 transponders for a hybrid system by using the benefit of propellant mass reduction. It takes 4-8 months to reach geostationary orbit for full electric satellite but in a chemical propulsion satellite, it takes a few days. Satellite subjects to additional radiation due to this long trip and hence additional aging. Expected revenue loss is another issue and conventional insurance policy needs an amendment to be in line with electric propulsion technology. The development of full electric satellite lowers the cost per transponder significantly.

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.

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.

THE PURPOSE. Consider the issue of the presence of distortions in the form of sinusoidal voltage and current in autonomous electrical power systems with electrical propulsion systems, built on the principle of the unified electric power system. Compare the results of the study of voltage distortion for electrical power systems with electrical propulsion systems of various structures and give recommendations for their application.METHODS. For research, unified electric power systems with DC electrical propulsion systems on the Yeysk ferry and alternating current on the asymmetric icebreaker Baltika are considered. The possibilities of programming with frequency control of modern drives of rudder propellers with AC motors are analyzed.RESULTS. Oscillograms of voltages and currents of generators were obtained using thyristor converters and inverters to control electrical propulsion systems in various modes. Significant ripples and distortions of the sinusoidal voltage and generator currents were noted when thyristor converters are used to power DC propulsion electric motors. Also, small deviations from the sinusoidal form of the ship's network voltage were recorded when inverters were used to control AC propulsion motors as part of rudder propellers.CONCLUSION. The issue of ensuring the quality of electricity produced in autonomous electrical power systems of sea vessels is of great relevance and importance. To ensure the best results, it is advisable to use azimuthing podded drive with AC electric motors, the rotation speed of which is implemented by inverter frequency converters with a DC link.

Electric, hybrid, and turboelectric fixed-wing aircraft: A review of concepts, models, and design approaches

Active development of electric thrustors began 10 years ago. Today, several kinds of thrustors have achieved efficiencies above 90 % and lifetimes of several thousand hours. The following article derives the basic theory of electric thrust production at constant exhaust velocity, and at variable exhaust velocity programmed for optimum vehicle performance. Electrothermal or arcjet; electrostatic or ion; and electrodynamic or plasma thrustors are described. At the present time, ion thrustors of the electron bombardment and of the surface ionization types are the most promising systems. Electric power in space may be generated by solar cells or nuclear-electric generators. It is expected that the incore thermionic converter will eventually be the preferred system. A variety of missions with electric propulsion systems appear feasible and highly desirable, among them orbital station keeping, attitude control, planetary probes, solar and out-of-the-ecliptic probes, deep-space probes, and manned Mars and Venus exploration. For each mission, a careful systems-design study must be made, which will provide the optimum selection of thrustor type, thrust level, exhaust velocity, thrust program, power source, trajectory, and flight plan.

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.

The electric propulsion system of an electric aircraft is the core power of the whole aircraft, which provides assurance for the safe flight of the aircraft. There are multiple stages in the flight process of electric aircraft, as well as different failure behaviors of system components at different stages. The electric motor system is a k-out-of-n system. Therefore, it is not possible to accurately describe the electric motor system with the existing operators when using the GO-FLOW method for electric propulsion system reliability calculations. In order to solve this problem, this paper models the electric motor system using the multi-valued decision diagram method (MDD method), based on which a new operator 43 is proposed to simulate the electric motor system, and then builds the GO-FLOW model and calculates the overall reliability based on the distributed electric propulsion system schematic. The example results show that the method can calculate the reliability of the distributed electric propulsion system more realistically.

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

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.

The mission targets of the Chinese asteroid exploration are an earth co-orbiting celestial body 2016HO3 and the main-belt comet in the asteroid belt. Compared with Lunar Exploration and Mars Exploration missions, asteroid exploration mission is special. The light intensity varies widely during the mission, resulting in a large change in the output power range of the solar array, which seriously affects the power system design and the load design during the mission; the spacecraft uses a high-power electric propulsion system; the power system has strict weight requirements, the existing power supply system design cannot meet the requirements. Aiming at the design constraints of the large variation of solar wing light intensity, the large difference in power requirements between the electric propulsion and the platform, and the lightweight design of the spacecraft, a dual-bus power supply system is proposed. High-voltage bus power the electric propulsion system. The solar cell array directly converges to form a high-voltage bus, which supplies power for the electric propulsion system. The low-voltage bus adopts a semi-adjustable bus, and the solar cell array adopts a circular flexible solar wing. The main detector and return cabin are equipped with lithium-ion battery pack and lithium fluorocarbon battery pack respectively. Aiming at the requirement of long-term on-orbit operation of electric propulsion system, a solar wing output power measurement system is designed. Product test results show that the power system design can meet the working requirements of electric propulsion and platform loads. The power system design can provide reference for the other deep space exploration power system.

Power processing units (PPUs) in an electric propulsion system provide many challenging integration issues. The PPU must provide power to the electric thruster while maintaining compatibility with all of the spacecraft power and data systems. Inefficiencies in the power processor produce heat, which must be radiated to the environment in order to ensure reliable operation. Although PPU efficiencies are generally greater than 0.9, heat loads are often substantial. This heat must be rejected by thermal control systems which generally have specific masses of 15-30 kg/kW. PPUs also represent a large fraction of the electric propulsion system dry mass. Simplification or elimination of power processing in a propulsion system would reduce the electric propulsion system specific mass and improve the overall reliability and performance. A direct drive system would eliminate all or some of the power supplies required to operate a thruster by directly connecting the various thruster loads to the solar array. The development of concentrator solar arrays has enabled power bus voltages in excess of 300 V which is high enough for direct drive applications for Hall thrusters such as the Stationary Plasma Thruster (SPT). The option of solar array direct drive for SPTs is explored to provide a comparison between conventional and direct drive system mass.

Electric propulsion systems derive their low overall mass, relative to chemical propulsion systems, from an optimized mix of independently controlled power and mass flow rate. A significant reduction in mass can be secured by substituting electric propulsion for conventional hydrazine systems for all the major propulsion tasks of a communications satellite. Additional advantages for all electric propulsion maneuvers are precision location and higher pointing accuracy. Three auxiliary electric propulsion systems are now available for application on communications satellites: the electrically augmented hydrazine system, the Teflon pulsed plasma system, and the mercury ion thruster system. Primary electric propulsion will be available in the mid to late 1980s as a spin-off of NASA's program of interplanetary exploration.

The use of fossil energy and coal has a significant impact on global warming through increased greenhouse gas emissions, particularly carbon dioxide (CO2). To address this issue, green technologies offer sustainable solutions, including the development of electric propulsion systems on unmanned aircraft (UAVs). Electric propulsion uses clean energy sources such as batteries and hydrogen fuel cells, significantly reducing emissions and air pollution. This research explores the importance of applying green technologies in the optimization of electric propulsion on UAVs, and the resulting environmental and economic benefits. Electric propulsion technology not only reduces carbon footprint and air pollution, but also improves energy efficiency and extends the operational life of UAVs. In addition, it allows UAVs to operate more quietly, reduces noise pollution, and can be used in various sensitive environments. The implementation of green technologies in UAV electric propulsion systems faces challenges such as limited battery capacity and recharging infrastructure. However, with collaboration between industry, government and research institutions, as well as regulatory support and incentives, the development of these technologies can be accelerated to achieve more environmentally friendly and sustainable UAV operations. Through the application of electric propulsion technology, UAVs are expected to provide innovative and sustainable solutions in areas such as aerial surveying, security surveillance, and freight forwarding, while making a positive impact on the environment and sustainability of the global ecosystem.